High power epicyclic gearbox and operation thereof

ABSTRACT

An engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox. The gearbox receives an input from a gearbox input shaft portion of the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox including a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The carrier and the gearbox input shaft each have a torsional stiffness, and a carrier to gearbox input shaft torsional stiffness ratio is greater than or equal to 70.

This application is a Continuation of U.S. Ser. No. 16/807,956 filedMar. 3, 2020, which is based upon and claims the benefit of priorityfrom GB 1917777.3 filed Dec. 5, 2019, the entire contents of the priorapplications being incorporated herein by reference.

The present disclosure relates to gearboxes for use in aircraft engines,to aircraft engines comprising such a gearbox, and to methods ofoperating such an aircraft. Such gearboxes may be epicyclic gearboxescomprising a planet carrier having stiffnesses meeting specifiedcriteria.

As used herein, a range “from value X to value Y” or “between value Xand value Y”, or the likes, denotes an inclusive range; including thebounding values of X and Y. As used herein, the term “axial plane”denotes a plane extending along the length of an engine, parallel to andcontaining an axial centreline of the engine, and the term “radialplane” denotes a plane extending perpendicular to the axial centrelineof the engine, so including all radial lines at the axial position ofthe radial plane. Axial planes may also be referred to as longitudinalplanes, as they extend along the length of the engine. A radial distanceor an axial distance is therefore a distance extending in a radialdirection in a radial plane, or a distance extending in an axialdirection in an axial plane, respectively.

According to a first aspect there is provided a gas turbine engine foran aircraft, the engine comprising an engine core comprising a turbine,a compressor, and a core shaft connecting the turbine to the compressor;a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and a gearbox arranged to receive an input fromthe core shaft and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft. The gearbox is anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted.The radial bending stiffness of the planet carrier is equal to orgreater than 1.20×10⁹ N/m.

The radial bending stiffness of the planet carrier may be less than orequal to 1.00×10¹² N/m.

The radial bending stiffness of the planet carrier may be equal to orgreater than 2.0×10⁹ N/m, and/or optionally in the range from 1.20×10⁹to 1.00×10¹² N/m or from 2.0×10⁹ to 1.5×10¹¹ N/m.

The tilt stiffness of the planet carrier may be greater than or equal to6.00×10⁸ Nm/radian, and optionally may be in the range from 1.3×10⁹ to1.2×10¹¹ Nm/radian.

The fan may have a fan diameter in the range from 240 to 280 cm. In suchembodiments, the radial bending stiffness of the planet carrier may beequal to or greater than 1.5×10⁹ N/m, and optionally less than or equalto 5×10¹⁰ N/m.

Alternatively, the fan may have a fan diameter in the range from 330 to380 cm. In such embodiments, the radial bending stiffness of the planetcarrier may be equal to or greater than 2.0×10⁹ N/m, and optionally lessthan or equal to 1.6×10¹¹ N/m.

According to a second aspect, there is provided a gas turbine engine foran aircraft, the engine comprising an engine core comprising a turbine,a compressor, and a core shaft connecting the turbine to the compressor;a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and a gearbox arranged to receive an input fromthe core shaft and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft. The gearbox is anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted.The tilt stiffness of the planet carrier is greater than or equal to6.0×10⁸ Nm/rad.

The tilt stiffness of the planet carrier may be less than or equal to2.8×10¹¹ Nm/rad.

Optionally, the tilt stiffness of the planet carrier may be greater thanor equal to 1.3×10⁹ Nm/radian, and optionally in the range from 1.3×10⁹to 1.2×10¹¹ Nm/radian.

The radial bending stiffness of the planet carrier may be equal to orgreater than 1.20×10⁹ N/m, and optionally in the range from 1.20×10⁹ to1×10¹² N/m or from 2.0×10⁹ to 1.5×10¹¹ N/m.

The fan may have a fan diameter in the range from 240 to 280 cm. In suchembodiments, the tilt stiffness of the planet carrier may be greaterthan or equal to 2.2×10⁹ Nm/radian, and optionally less than or equal to1.4×10¹¹ Nm/radian.

Alternatively, the fan may have a fan diameter in the range from 330 to380 cm. In such embodiments, the tilt stiffness of the planet carriermay be greater than or equal to 2.3×10⁹ Nm/radian, and optionally lessthan or equal to 2.8×10¹¹ Nm/radian.

According to a third aspect, there is provided a gas turbine engine foran aircraft, the engine comprising an engine core comprising a turbine,a compressor, and a core shaft connecting the turbine to the compressor;a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and a gearbox arranged to receive an input fromthe core shaft and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft. The gearbox is anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted.The torsional stiffness of the planet carrier is greater than or equalto 1.60×10⁸ Nm/rad, and optionally smaller than or equal to 1.00×10¹¹Nm/rad.

The inventor has discovered that maintaining one or more stiffnesses ofthe carrier in the ranges claimed herein may allow for compensation ingear misalignment (for example due to variation within manufacturingtolerances and/or wear during operation) whilst avoiding significantdistortion of the gearbox. The compensation may result in reducedvariability in the loading across and/or between gears (e.g. a more evenload share). In turn this may allow the mass of the gears to be reducedwhilst still retaining the required life and efficiency for an aircraftapplication

One or more stiffnesses of the carrier may therefore be selected to berelatively high to reduce or avoid deleterious wind-up or distortion ofthe carrier and/or misalignment of the gears carried by the carrier. Itmay be beneficial for the one or more of the stiffnesses to remain lowenough to allow sufficient flexibility to correct for any minor gearmisalignment due to manufacturing issues. The inventor discovered thatin some arrangements maintaining one or more of the stiffnesses withinthe corresponding specified ranges provides a beneficial combination ofthese effects.

It has been found that maintaining an even distribution of load betweenthe planet gears is desirable to improve gearbox lifetime andreliability. The inventor discovered that maintaining one or more of thecarrier stiffnesses within the applicable specified range allowssufficient flexibility of the carrier to facilitate obtaining a moreeven load share (i.e. an improved load-share factor) by allowing theplanet gears to move relative to each other and/or relative to thecarrier. In some arrangements, if the carrier stiffness were too high,load-share factor may deteriorate due to an inability to accommodate anypre-existing misalignment, or any misalignment arising during use.

The carrier design of various embodiments, with one or more stiffnessesas defined, may therefore facilitate obtaining and/or maintainingcorrect gear alignment.

One or more of the following features may apply to any of the precedingthree aspects:

The torsional stiffness of the planet carrier may be greater than orequal to 1.60×10⁸ Nm/rad, and optionally in the range from 1.60×10⁸ to1.00×10¹¹ Nm/rad, or from 2.7×10⁸ to 1×10¹⁰ Nm/rad.

A pitch circle diameter of pins on which the planet gears are mountedmay be in the range from 0.38 to 0.65 m, and optionally may be equal to0.4 m or 0.55 m.

According to a fourth aspect, there is provided a method of operation ofa gas turbine engine for an aircraft comprising an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; and a gearbox arranged toreceive an input from the core shaft and to output drive to the fan soas to drive the fan at a lower rotational speed than the core shaft, thegearbox being an epicyclic gearbox comprising a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier on which the planetgears are mounted, and wherein the radial bending stiffness of theplanet carrier is equal to or greater than 1.20×10⁹ N/m, and optionallyless than or equal to 1.00×10¹² N/m. The method comprises operating thegas turbine engine to provide propulsion under cruise conditions.

According to a fifth aspect, there is provided a method of operation ofa gas turbine engine for an aircraft comprising an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; and a gearbox arranged toreceive an input from the core shaft and to output drive to the fan soas to drive the fan at a lower rotational speed than the core shaft, thegearbox being an epicyclic gearbox comprising a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier on which the planetgears are mounted, and wherein the tilt stiffness of the planet carrieris greater than or equal to 6.0×10⁸ Nm/rad. Additionally oralternatively, in a further aspect, the torsional stiffness of theplanet carrier may be greater than or equal to 1.60×10⁸ Nm/rad, andoptionally less than or equal to 1.00×10¹¹ Nm/rad. The method comprisesoperating the gas turbine engine to provide propulsion under cruiseconditions.

The method of the fourth or fifth aspect may comprise driving thegearbox with an input torque of:

-   -   (i) greater than or equal to 10,000 Nm, and optionally of 10,000        to 50,000 Nm at cruise; and/or    -   (ii) greater than or equal to 28,000 Nm, and optionally of        28,000 to 135,000 Nm at MTO.

According to a sixth aspect, there is provided a propulsor for anaircraft, the propulsor comprising a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is an epicyclic gearbox arranged to receive an input from acore shaft driven by the power unit and to output drive to the fan so asto drive the fan at a lower rotational speed than the core shaft, andcomprises a sun gear, a plurality of planet gears, a ring gear, and aplanet carrier on which the planet gears are mounted. The radial bendingstiffness of the planet carrier is equal to or greater than 1.20×10⁹N/m. Optionally, the radial bending stiffness of the planet carrier maybe less than or equal to 1.00×10¹² N/m.

According to a seventh aspect, there is provided a propulsor for anaircraft, the propulsor comprising a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is an epicyclic gearbox arranged to receive an input fromthe core shaft and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft, and comprises a sun gear,a plurality of planet gears, a ring gear, and a planet carrier on whichthe planet gears are mounted. The tilt stiffness of the planet carrieris greater than or equal to 6.00×10⁸ Nm/rad. The tilt stiffness of theplanet carrier may be less than or equal to 2.8×10¹¹ Nm/rad.

Additionally or alternatively, in a further aspect, the torsionalstiffness of the planet carrier may be greater than or equal to 1.60×10⁸Nm/rad, and optionally smaller than or equal to 1.00×10¹¹ Nm/rad.

The propulsor of the sixth or seventh aspect may have any or all of thefeatures as described for the gas turbine engine of the first, second,and/or third aspect.

According to an eighth aspect there is provided a gas turbine engine foran aircraft, the engine comprising an engine core comprising a turbine,a compressor, and a core shaft connecting the turbine to the compressor;a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and a gearbox arranged to receive an input fromthe core shaft and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft. The gearbox is anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted. Aradial bending to torsional carrier stiffness ratio of:

$\frac{{the}\mspace{14mu}{radial}\mspace{14mu}{bending}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{effective}\mspace{14mu}{linear}\mspace{14mu}{torsional}{\mspace{11mu}\;}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}$

-   -   is greater than or equal to 0.030.

The radial bending to torsional carrier stiffness ratio may be less thanor equal to 2.0×10⁰ (i.e. 2.0).

The radial bending to torsional carrier stiffness ratio may be in therange from in the range from 0.030 to 2.0, and optionally in the rangefrom 0.060 to 1.0.

The radial bending stiffness of the planet carrier may be equal to orgreater than 1.20×10⁹ N/m, and optionally in the range from 1.20×10⁹ to1.00×10¹² N/m or from 2.0×10⁹ to 1.5×10¹¹ N/m.

The effective linear torsional stiffness of the planet carrier may begreater than or equal to 7.00×10⁹ N/m, and optionally in the range from7.00×10⁹ to 1.20×10¹¹ N/m or from 9.1×10⁹ to 8.0×10¹⁰ N/m.

The tilt stiffness of the planet carrier may be greater than or equal to6.00×10⁸ Nm/radian, and optionally in the range from 1.3×10⁹ to 1.2×10¹¹Nm/radian.

The radial bending to torsional carrier stiffness ratio may be in therange from 0.060 to 0.30. Alternatively, the radial bending to torsionalcarrier stiffness ratio may be in the range from 0.30 to 2.0.

A tilt to torsional carrier stiffness ratio of:

$\frac{{the}\mspace{14mu}{tilt}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{{to}{rsional}}{\mspace{11mu}\;}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}$

-   -   may be in the range from 0.7 to 20.

According to a ninth aspect, there is provided a gas turbine engine foran aircraft, the engine comprising an engine core comprising a turbine,a compressor, and a core shaft connecting the turbine to the compressor;a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and a gearbox arranged to receive an input fromthe core shaft and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft. The gearbox is anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted. Atilt to torsional carrier stiffness ratio of:

$\frac{{the}\mspace{14mu}{tilt}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{{to}{rsional}}{\mspace{11mu}\;}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}$

-   -   is in the range from 0.7 to 20.

The tilt to torsional carrier stiffness ratio may be in the range from0.7 to 7.3.

The tilt stiffness of the planet carrier may be greater than or equal to6.00×10⁸ Nm/radian, and optionally in the range from 1.3×10⁹ to 1.2×10¹¹Nm/radian.

The radial bending stiffness of the planet carrier may be equal to orgreater than 1.20×10⁹ N/m, and optionally in the range from 1.20×10⁹ to1×10¹² N/m or from 2.0×10⁹ to 1.5×10¹¹ N/m.

The torsional stiffness of the planet carrier may be greater than orequal to 1.60×10⁸ Nm/radian, and optionally in the range from 1.60×10⁸to 1.00×10¹¹ Nm/radian, or from 2.7×10⁸ to 1×10¹⁰ Nm/radian.

The fan may have a fan diameter in the range from 240 to 280 cm. In suchembodiments, the tilt to torsional carrier stiffness ratio may be in therange from 2.5 to 8.0.

Alternatively, the fan may have a fan diameter in the range from 330 to380 cm. In such embodiments, the tilt to torsional carrier stiffnessratio may be in the range from 1.5 to 7.9.

A radial bending to torsional carrier stiffness ratio of:

$\frac{\text{the radial bending stiffness of the planet carrier}}{\text{the effective linear torsional stiffness of the planet carrier}}$

-   -   may be in the range from 0.030 to 2.0.

The inventor discovered that maintaining the ratio of the radial bendingor tilt stiffness of the carrier to the torsional stiffness of thecarrier within the specified range allows for improved avoidance ofdamage to gear teeth (due to a relative increase in the carriertorsional stiffness). Any further relative increase of the torsionalstiffness was found not to provide further benefits in tooth protection,and may instead risk reducing overall performance by adding unnecessarysize and/or weight to the carrier. The torsional stiffness of thecarrier is therefore arranged to be relatively high to reduce or avoidthe risk of distortion of gear teeth. In particular, the inventorappreciated that wind-up of the carrier whilst the gear teeth are meshedmay chip, grind or deform gear teeth as they are forced against teeth ofthe opposing gear.

According to a tenth aspect, there is provided a method of operation ofa gas turbine engine for an aircraft, the engine comprising an enginecore comprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; and a gearbox arranged toreceive an input from the core shaft and to output drive to the fan soas to drive the fan at a lower rotational speed than the core shaft. Thegearbox is an epicyclic gearbox comprising a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier on which the planetgears are mounted. A radial bending to torsional carrier stiffness ratioof:

$\frac{\text{the radial bending stiffness of the planet carrier}}{\text{the effective linear torsional stiffness of the planet carrier}}$is in the range from 0.030 to 2.0. Additionally or alternatively, a tiltto torsional carrier stiffness ratio of:

$\frac{\text{the tilt stiffness of the planet carrier}}{\text{the torsional stiffness of the planet carrier}}$is in the range from 0.7 to 20.

The method comprises operating the gas turbine engine to providepropulsion under cruise conditions.

The method may comprise driving the gearbox with an input torque ofgreater than or equal to 10,000 Nm, and optionally of 10,000 to 50,000Nm at cruise.

The method may comprise driving the gearbox with an input torque ofgreater than or equal to 28,000 Nm, and optionally of 28,000 to 135,000Nm at MTO.

The engine used for the method of the tenth aspect may have any or allof the features of the engine of the eighth or ninth aspects.

According to an eleventh aspect, there is provided a propulsor for anaircraft, the propulsor comprising a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is an epicyclic gearbox arranged to receive an input fromthe core shaft and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft, and comprises a sun gear,a plurality of planet gears, a ring gear, and a planet carrier on whichthe planet gears are mounted. A radial bending to torsional carrierstiffness ratio of:

$\frac{\text{the radial bending stiffness of the planet carrier}}{\text{the effective linear torsional stiffness of the planet carrier}}$is in the range from 0.030 to 2.0.

According to a twelfth aspect, there is provided a propulsor for anaircraft, the propulsor comprising a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is an epicyclic gearbox arranged to receive an input fromthe core shaft and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft, and comprises a sun gear,a plurality of planet gears, a ring gear, and a planet carrier on whichthe planet gears are mounted. A tilt to torsional carrier stiffnessratio of:

$\frac{\text{the tilt stiffness of the planet carrier}}{\text{the torsional stiffness of the planet carrier}}$is in the range from 0.7 to 20.

The propulsor of the eleventh or twelfth aspect may have any or all ofthe features of the engine of the eighth or ninth aspects.

According to a thirteenth aspect there is provided a gas turbine enginefor an aircraft, the engine comprising an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a gearbox that receives aninput from a gearbox input shaft portion of the core shaft and outputsdrive to the fan so as to drive the fan at a lower rotational speed thanthe core shaft, the gearbox being an epicyclic gearbox comprising a sungear, a plurality of planet gears, a ring gear, and a planet carrier onwhich the planet gears are mounted. A carrier to gearbox input shafttorsional stiffness ratio of:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the gearbox input shaft}}$

-   -   is greater than or equal to 70, and optionally less than or        equal to 5.0×10³.

The inventor has discovered that the torsional stiffness of a gearboxsystem—including in particular the gearbox input shaft and thecarrier—should be distributed in the claimed proportion because keepingthe torsional stiffness of the carrier relatively high was found toreduce or avoid the risk of distortion of gear teeth as described above,but relatively increasing the torsional stiffness of the core shaft (theinput shaft to the gearbox) did not improve this effect, but ratherdeleteriously added size and/or weight to the shaft without acorresponding benefit.

Whilst torsional flexibility of the gearbox input shaft may have less ofan effect on gearbox performance than carrier flexibility, the skilledperson would appreciate that too low a gearbox input shaft torsionalstiffness may result in wind-up of the sun gear, potentially resultingin gear misalignments. Relative increases in gearbox input shaftstiffness beyond the claimed ratio range may provide no further benefit,and may instead unnecessarily increase size and/or weight of the fanshaft.

The carrier to gearbox input shaft torsional stiffness ratio may beequal to or greater than 75, and optionally in the range from 7.5×10¹ to3×10³.

The torsional stiffness of the planet carrier may be greater than orequal to 1.60×10⁸ Nm/rad, and optionally in the range from 1.60×10⁸ to1.00×10¹¹ Nm/rad, or from 2.7×10⁸ to 1×10¹⁰ Nm/rad

The torsional stiffness of the gearbox input shaft may be equal to orgreater than 1.4×10⁶ Nm/radian, and optionally in the range from 1.4×10⁶to 2.5×10⁸ Nm/radian.

The fan may have a fan diameter in the range from 240 to 280 cm. In suchembodiments, the carrier to gearbox input shaft torsional stiffnessratio may be greater than or equal to 7.3×10¹, and optionally less thanor equal to 1.0×10³.

The fan may have a fan diameter in the range from 330 to 380 cm. In suchembodiments, the carrier to gearbox input shaft torsional stiffnessratio may be greater than or equal to 1.0×10², and optionally less thanor equal to 5.0×10³.

The gas turbine engine may further comprise a gearbox support arrangedto support the gearbox in a fixed position within the engine and havinga torsional stiffness. A carrier to gearbox support torsional stiffnessratio of:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the gearbox support}}$

-   -   may be greater than or equal to 2.3, and optionally greater than        or equal to 2.6.

The gas turbine engine may further comprise a fan shaft that connects anoutput of the gearbox to the fan. A carrier to fan shaft stiffness ratioof:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the fan shaft}}$

-   -   may be greater than or equal to 8, and optionally greater than        or equal to 9.

The product of the torsional stiffness of the planet carrier and thetorsional stiffness of the gearbox input shaft may be greater than orequal to 1.5×10¹⁴ N²m²rad⁻², and optionally greater than or equal to2.2×10¹⁴ N²m²rad⁻².

The fan may have a fan diameter in the range from 240 to 280 cm and theproduct of the torsional stiffness of the planet carrier and thetorsional stiffness of the gearbox input shaft may be greater than orequal to 1.5×10¹⁴ N² m²rad⁻².

The fan may have a fan diameter in the range from 330 to 380 cm and theproduct of the torsional stiffness of the planet carrier and thetorsional stiffness of the gearbox input shaft may be greater than orequal to 3.0×10¹⁵ N² m²rad⁻².

The gearbox may be a star gearbox, in which the planet carrier does notrotate in use.

A pitch circle diameter of pins on which the planet gears are mountedmay be in the range from 0.38 to 0.65 m, and optionally may be equal to0.4 m or 0.55 m.

According to a fourteenth aspect, there is provided a method ofoperation of a gas turbine engine for an aircraft comprising: an enginecore comprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; and a gearbox arranged toreceive an input from a gearbox input shaft portion of the core shaftand to output drive to the fan so as to drive the fan at a lowerrotational speed than the core shaft, the gearbox being an epicyclicgearbox comprising a sun gear, a plurality of planet gears, a ring gear,and a planet carrier on which the planet gears are mounted.

A carrier to gearbox input shaft torsional stiffness ratio of:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the gearbox input shaft}}$is greater than or equal to 70. The method comprises operating the gasturbine engine to provide propulsion under cruise conditions. Thecarrier to gearbox input shaft torsional stiffness ratio may be lessthan or equal to 5,000.

The method may comprise driving the gearbox with an input torque ofgreater than or equal to 10,000 Nm, and optionally of 10,000 to 50,000Nm at cruise.

The method may comprise driving the gearbox with an input torque ofgreater than or equal to 28,000 Nm, and optionally of 28,000 to 135,000Nm at MTO.

According to a fifteenth aspect, there is provided a propulsor for anaircraft, the propulsor comprising: a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is an epicyclic gearbox arranged to receive an input from agearbox input shaft portion of a core shaft driven by the power unit andto output drive to the fan so as to drive the fan at a lower rotationalspeed than the core shaft, and comprises a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier on which the planetgears are mounted. A carrier to gearbox input shaft torsional stiffnessratio of:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the gearbox input shaft}}$is greater than or equal to 70. The carrier to gearbox input shafttorsional stiffness ratio may be less than or equal to 5,000.

The propulsor may have some or all of the features described above withrespect to the gas turbine engine, and may be a gas turbine engine insome embodiments.

In various other aspects of the invention, the specified boundaries onthe carrier to gearbox input shaft torsional stiffness ratio may bereplaced by, or provided in addition to, specified boundaries on theproduct of the components of the carrier to gearbox input shafttorsional stiffness ratio, i.e. boundaries on the torsional stiffness ofthe planet carrier multiplied by the torsional stiffness of the gearboxinput shaft. The value of this product, in various aspects, may begreater than or equal to 1.5×10¹⁴ N²m²rad⁻², and optionally less than1.0×10¹⁷ N²m²rad⁻². Optionally, the value may be greater than or equalto 2.2×10¹⁴ N²m²rad⁻², and optionally less than 5.0=10¹⁶ N²m²rad⁻².

For example, according to another aspect, there is provided a gasturbine engine for an aircraft comprising: an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a gearbox that receives aninput from a gearbox input shaft portion of the core shaft and outputsdrive to the fan so as to drive the fan at a lower rotational speed thanthe core shaft, the gearbox being an epicyclic gearbox comprising a sungear, a plurality of planet gears, a ring gear, and a planet carrier onwhich the planet gears are mounted. The torsional stiffness of theplanet carrier multiplied by the torsional stiffness of the gearboxinput shaft is greater than or equal to 1.5×10¹⁴ N² m²rad⁻².

The skilled person would appreciate that method and propulsor aspectsmay be formulated accordingly. The optional features of the aspects forthe corresponding ratio may also apply to these aspects.

According to a sixteenth aspect, there is provided a gas turbine enginefor an aircraft, the engine comprising: an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft, the gearbox being anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted;and a gearbox support arranged to support the gearbox in a fixedposition within the engine and having a torsional stiffness. A carrierto gearbox support torsional stiffness ratio of:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the gearbox support}}$

-   -   is greater than or equal to 2.3. The carrier to gearbox support        torsional stiffness ratio may be less than or equal to 300.

The inventor has discovered that the torsional stiffness of a gearboxsystem—including in particular the gearbox support and thecarrier—should be distributed in the claimed proportion to provideimprovements such as those described above with respect to thethirteenth aspect. In particular, the inventor discovered thatmaintaining the ratio within the specified range allowed for a reducedrisk of gear tooth damage whilst still maintaining sufficient stiffnessto avoid damagingly large amplitudes of torsional vibration modes of thegearbox support.

The carrier to gearbox support torsional stiffness ratio may be greaterthan or equal to 2.6, and optionally in the range from 2.6 to 50.

The torsional stiffness of the planet carrier may be greater than orequal to 1.60×10⁸ Nm/rad, and optionally in the range from 1.60×10⁸ to1.00×10¹¹ Nm/rad, or from 2.7×10⁸ to 1×10¹⁰ Nm/rad.

The torsional stiffness of the gearbox input shaft portion of the coreshaft may be greater than or equal to 1.4×10⁶ Nm/radian, and optionallygreater than or equal to 1.6×10⁶ Nm/radian.

The fan may have a fan diameter in the range from 240 to 280 cm, and thecarrier to gearbox support torsional stiffness ratio may be greater thanor equal to 2.3.

The fan may have a fan diameter in the range from 330 to 380 cm, and thecarrier to gearbox support torsional stiffness ratio may be greater thanor equal to 3.5.

The torsional stiffness of the planet carrier multiplied by thetorsional stiffness of the gearbox may be greater than or equal to5×10¹⁵ N²m²rad⁻², and optionally less than 1.0×10¹⁹ N²m²rad⁻².

The core shaft may comprise a gearbox input shaft portion arranged toprovide the input to the gearbox. A carrier to gearbox input shafttorsional stiffness ratio of:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the gearbox input shaft}}$may be greater than or equal to 70.

The carrier to gearbox input shaft torsional stiffness ratio may beequal to or greater than 75, and optionally in the range from 7.5×10¹ to3×10³.

The gearbox may be a star gearbox, in which the planet carrier does notrotate in use.

A pitch circle diameter of pins on which the planet gears are mountedmay be in the range from 0.38 to 0.65 m, and optionally may be equal to0.4 m or 0.55 m.

The gas turbine engine may further comprise a fan shaft that connects anoutput of the gearbox to the fan. A carrier to fan shaft stiffness ratioof:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the fan shaft}}$may be greater than or equal to 8, and optionally greater than or equalto 9, and may be less than or equal to 1,100.

According to a seventeenth aspect, there is provided a method ofoperation of a gas turbine engine for an aircraft, the engine comprisingan engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; a fan located upstream of theengine core, the fan comprising a plurality of fan blades; and a gearboxthat receives an input from the core shaft and outputs drive to the fanso as to drive the fan at a lower rotational speed than the core shaft.The gearbox is an epicyclic gearbox comprising a sun gear, a pluralityof planet gears, a ring gear, and a planet carrier on which the planetgears are mounted; and a gearbox support arranged to support the gearboxin a fixed position within the engine and having a torsional stiffness.A carrier to gearbox support torsional stiffness ratio of:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the gearbox support}}$

-   -   is greater than or equal to 2.3. The method comprises operating        the gas turbine engine to provide propulsion under cruise        conditions. The carrier to gearbox support torsional stiffness        ratio may be less than or equal to 300.

The method may further comprise driving the gearbox with an input torqueof greater than or equal to 10,000 Nm, and optionally of 10,000 to50,000 Nm at cruise.

The method may further comprise driving the gearbox with an input torqueof greater than or equal to 28,000 Nm, and optionally of 28,000 to135,000 Nm at MTO.

According to an eighteenth aspect, there is provided a propulsor for anaircraft, the propulsor comprising: a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is arranged to receive an input from a core shaft driven bythe power unit and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft, the gearbox being anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted.The propulsor further comprises a gearbox support arranged to supportthe gearbox in a fixed position within the propulsor and having atorsional stiffness. A carrier to gearbox support torsional stiffnessratio of:

$\frac{\text{the}\text{torsional}\text{stiffness of the planet carrier}}{\text{the torsional stiffness of the gearbox support}}$

-   -   is greater than or equal to 2.3, and optionally less than or        equal to 300.

The propulsor may have some or all of the features described above withrespect to the gas turbine engine, and may be a gas turbine engine insome embodiments.

In various other aspects of the invention, the specified boundaries onthe carrier to gearbox support torsional stiffness ratio may be replacedby, or provided in addition to, specified boundaries on the product ofthe components of the carrier to gearbox support torsional stiffnessratio, i.e. boundaries on the torsional stiffness of the planet carriermultiplied by the torsional stiffness of the gearbox support. The valueof this product, in various aspects, may be greater than or equal to5×10¹⁵ N² m²rad⁻², and optionally less than 1.0×10¹⁹ N² m²rad⁻².Optionally, the value may be greater than or equal to 8.0×10¹⁵N²m²rad⁻², and optionally less than 2.0×10¹⁸ N² m²rad⁻².

For example, according to another aspect, there is provided a gasturbine engine for an aircraft, the engine comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft, the gearboxbeing an epicyclic gearbox comprising a sun gear, a plurality of planetgears, a ring gear, and a planet carrier on which the planet gears aremounted; and a gearbox support arranged to support the gearbox in afixed position within the engine and having a torsional stiffness. Atorsional stiffness of the planet carrier multiplied by a torsionalstiffness of the gearbox support is greater than or equal to 5×10¹⁵ N²m²rad⁻².

The skilled person would appreciate that method and propulsor aspectsmay be formulated accordingly. The optional features of the aspects forthe corresponding ratio may also apply to these aspects.

According to a nineteenth aspect, there is provided a gas turbine enginefor an aircraft, the engine comprising: an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft, the gearbox being anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted;and a fan shaft that connects an output of the gearbox to the fan. Acarrier to fan shaft stiffness ratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}{\mspace{11mu}\;}{carrier}}{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{fan}{\mspace{11mu}\;}{shaft}}$

-   -   is greater than or equal to 8. The carrier to fan shaft        stiffness ratio may be less than or equal to 1,100.

The inventor has discovered that the torsional stiffness of a gearboxsystem—including in particular the carrier and the gearbox output shaft(fan shaft)—should be distributed in the claimed proportion to provideimprovements such as those described above with respect to thethirteenth and sixteenth aspects. In particular, the inventor discoveredthat maintaining the ratio within the specified range allowed for areduced risk of gear tooth damage whilst still maintaining sufficientstiffness of the fan shaft to avoid damagingly large displacements ofgears within the gearbox.

The skilled person would appreciate that too low a fan shaft stiffnessmay result in wind-up of the carrier (for a planetary gearbox) or of thering gear (for a star gearbox), resulting in gear misalignments.Relative increases in fan shaft stiffness beyond the claimed ratio rangemay provide no further benefit, however, and may instead deleteriouslyincrease size and/or weight of the fan shaft.

The carrier to fan shaft torsional stiffness ratio may be greater thanor equal to 9, and optionally in the range from 9 to 1.9×10².

The fan shaft may comprise two shaft portions; a gearbox output shaftportion extending from the gearbox and a fan portion extending betweenthe gearbox output shaft portion and the fan.

The torsional stiffness of the planet carrier may be greater than orequal to 1.60×10⁸ Nm/rad, and optionally in the range from 1.60×10⁸ to1.00×10¹¹ Nm/rad, or from 2.7×10⁸ to 1×10¹⁰ Nm/radian.

The torsional stiffness of the gearbox input shaft portion of the coreshaft may be greater than or equal to 1.4×10⁶ Nm/radian, and optionallygreater than or equal to 1.6×10⁶ Nm/radian.

The fan may have a fan diameter in the range from 240 to 280 cm, and thecarrier to fan shaft torsional stiffness ratio may be greater than orequal to 9.

The fan may have a fan diameter in the range from 330 to 380 cm, and thecarrier to fan shaft torsional stiffness ratio may be greater than orequal to 12.

The gearbox may be a star gearbox, in which the planet carrier does notrotate in use.

A pitch circle diameter of pins on which the planet gears are mountedmay be in the range from 0.38 to 0.65 m, and optionally may be equal to0.4 m or 0.55 m.

The gas turbine engine may further comprise a gearbox support arrangedto support the gearbox in a fixed position within the engine and havinga torsional stiffness. A carrier to gearbox support torsional stiffnessratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{gearbox}\mspace{14mu}{support}}$

-   -   may be greater than or equal to 2.3, and optionally greater than        or equal to 2.6.

The carrier to gearbox support torsional stiffness ratio may be in therange from 2.3 to 300, and optionally from 2.6 to 50.

The core shaft may comprise a gearbox input shaft portion arranged toprovide the input to the gearbox. A carrier to gearbox input shafttorsional stiffness ratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{torsional}{\mspace{11mu}\;}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{gearbox}\mspace{14mu}{input}\mspace{14mu}{shaft}}$

-   -   may be greater than or equal to 70, and optionally less than or        equal to 5,000.

The carrier to gearbox input shaft torsional stiffness ratio may beequal to or greater than 75, and optionally in the range from 7.5×10¹ to3×10³.

According to a twentieth aspect, there is provided a method of operationof a gas turbine engine for an aircraft, the engine comprising: anengine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; a fan located upstream of theengine core, the fan comprising a plurality of fan blades. The enginecore further comprises a gearbox that receives an input from the coreshaft and outputs drive to the fan so as to drive the fan at a lowerrotational speed than the core shaft, the gearbox being an epicyclicgearbox comprising a sun gear, a plurality of planet gears, a ring gear,and a planet carrier on which the planet gears are mounted; and a fanshaft that connects an output of the gearbox to the fan. A carrier tofan shaft stiffness ratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{torsional}{\mspace{11mu}\;}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{shaft}}$

-   -   is greater than or equal to 8, and optionally may be less than        or equal to 1,100.

The method comprises operating the gas turbine engine to providepropulsion under cruise conditions.

The method may further comprise driving the gearbox with an input torqueof greater than or equal to 10,000 Nm, and optionally of 10,000 to50,000 Nm at cruise.

The method may further comprise driving the gearbox with an input torqueof greater than or equal to 28,000 Nm, and optionally of 28,000 to135,000 Nm at MTO.

The engine may be as described in the nineteenth aspect.

According to a twenty-first aspect, there is provided a propulsor for anaircraft, the propulsor comprising: a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is arranged to receive an input from a core shaft driven bythe power unit and to output drive to the fan so as to drive the fan ata lower rotational speed than the core shaft, the gearbox being anepicyclic gearbox comprising a sun gear, a plurality of planet gears, aring gear, and a planet carrier on which the planet gears are mounted;and a fan shaft that connects an output of the gearbox to the fan. Acarrier to fan shaft stiffness ratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{torsional}{\mspace{11mu}\;}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{shaft}}$

-   -   is greater than or equal to 8, and optionally may be less than        or equal to 1,100.

The propulsor may have some or all of the features described above withrespect to the gas turbine engine, and may be a gas turbine engine insome embodiments.

In various other aspects of the invention, the specified boundaries onthe carrier to fan shaft stiffness ratio may be replaced by, or providedin addition to, specified boundaries on the product of the components ofthe carrier to fan shaft stiffness ratio, i.e. boundaries on thetorsional stiffness of the planet carrier multiplied by the torsionalstiffness of the fan shaft. The value of this product, in variousaspects, may be greater than or equal to 1.5×10¹⁵ N² m²rad⁻², andoptionally less than 3.0×10¹⁸ N² m²rad⁻². Optionally, the value may begreater than or equal to 2.0×10¹⁵ N² m²rad⁻², and optionally less than7.0×10¹⁷ N² m²rad⁻².

For example, according to another aspect, there is provided a gasturbine engine for an aircraft, the engine comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft, the gearboxbeing an epicyclic gearbox comprising a sun gear, a plurality of planetgears, a ring gear, and a planet carrier on which the planet gears aremounted; and a fan shaft that connects an output of the gearbox to thefan. The torsional stiffness of the planet carrier multiplied by thetorsional stiffness of the fan shaft is greater than or equal to1.5×10¹⁵ N² m²rad⁻².

The skilled person would appreciate that method and propulsor aspectsmay be formulated accordingly. The optional features of the aspects forthe corresponding ratio may also apply to these aspects.

According to a twenty-second aspect, there is provided a gas turbineengine for an aircraft, the engine comprising an engine core comprisinga turbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a gearbox arranged to receivean input from the core shaft and to output drive to the fan so as todrive the fan at a lower rotational speed than the core shaft. Thegearbox is an epicyclic gearbox comprising a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier comprising a pluralityof pins, each pin being arranged to have a planet gear of the pluralityof planet gears mounted thereon. A first carrier to pin stiffness ratioof:

$\frac{{the}\mspace{14mu}{effective}\mspace{14mu}{linear}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{radial}\mspace{14mu}{bending}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}}$

-   -   is greater than or equal to 10. The first carrier to pin        stiffness ratio may be less than or equal to 40.

The first carrier to pin stiffness ratio may be greater than or equal to15, and optionally in the range from 15 to 30.

The effective linear torsional stiffness of the planet carrier may begreater than or equal to 7.00×10⁹ N/m, and optionally in the range from7.00×10⁹ to 1.20×10¹¹ N/m or from 9.1×10⁹ to 8.0×10¹⁰ N/m.

The radial bending stiffness of each pin may be greater than or equal to3.00×10⁸ N/m, and optionally greater than or equal to 6.3×10⁸ N/m.

A second carrier to pin stiffness ratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{tilt}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}}$is greater than or equal to 24, and optionally greater than or equal to34, and optionally less than or equal to 180.

The fan may have a fan diameter in the range from 240 to 280 cm. In suchembodiments, the first carrier to pin stiffness ratio may be greaterthan or equal to 15, and optionally in the range from 15 to 25.Alternatively, the fan may have a fan diameter in the range from 330 to380 cm. In such embodiments, the first carrier to pin stiffness ratiomay be greater than or equal to 16, and optionally in the range from 16to 35.

According to a twenty-third aspect, there is provided a gas turbineengine for an aircraft, the engine comprising an engine core comprisinga turbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a gearbox arranged to receivean input from the core shaft and to output drive to the fan so as todrive the fan at a lower rotational speed than the core shaft. Thegearbox is an epicyclic gearbox comprising a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier comprising a pluralityof pins, each pin being arranged to have a planet gear of the pluralityof planet gears mounted thereon. A second carrier to pin stiffness ratioof:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{{the}\mspace{14mu}{tilt}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}}$

-   -   is greater than or equal to 24, and optionally may be less than        or equal to 180.

The second carrier to pin stiffness ratio may be greater than or equalto 34, and optionally in the range from 34 to 140.

The torsional stiffness of the planet carrier may be greater than orequal to 1.60×10⁸ Nm/rad, and optionally may be in the range from1.60×10⁸ to 1.00×10¹¹ Nm/rad, or from 2.7×10⁸ to 1×10¹⁰ Nm/rad.

The tilt stiffness of each pin may be greater than or equal to 4.0×10⁶Nm/rad, and optionally may be greater than or equal to 8.7×10⁶ Nm/rad.

A first carrier to pin stiffness ratio of:

$\frac{{{the}\mspace{14mu}{effective}\mspace{14mu}{linear}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}\mspace{14mu}}{{the}\mspace{14mu}{radial}\mspace{14mu}{bending}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}}$

-   -   may be greater than or equal to 10 or 15.

The first carrier to pin stiffness ratio may be less than or equal to40, and optionally in the range from 15 to 30.

The fan may have a fan diameter in the range from 240 to 280 cm. In suchembodiments, the second carrier to pin stiffness ratio may be greaterthan or equal to 34, and optionally in the range from 34 to 120.

Alternatively, the fan may have a fan diameter in the range from 330 to380 cm. In such embodiments, the second carrier to pin stiffness ratiomay be greater than or equal to 40, and optionally in the range from 40to 180.

The inventor has discovered that the torsional stiffness of the carrierand the radial bending and/or tilt stiffness of each pin of the carriershould be selected to match the claimed relationship so as to improveengine longevity and/or efficiency, for example by protecting gear teethand/or improving planet load sharing.

The torsional stiffness of the carrier is therefore arranged to berelatively high as compared to the radial bending and/or tilt stiffnessof each individual pin, to reduce or avoid the risk of distortion ofgear teeth as described above whilst also reducing the differentialload/improving load-share.

Further, the inventor discovered that the pin tilt stiffness may have amore significant effect than the pin radial bending stiffness—excesstilt deflections of the pin may be more damaging than radial bendingdeflections for the same magnitude of deflection as tilt deflections mayproduce two compounding effects—firstly, load share may worsen, withsome planet gears taking a larger share of the load than others, andsecondly face distribution of that load shifts. The larger force on aparticular planet gear is therefore concentrated on one side of the gearrather than equally distributed across the tooth. The increased load onthat gear and the increased concentration of that load may thereforedamage the gear teeth. Maintaining the pin tilt stiffness above 4.0×10⁶Nm/rad, and optionally above 8.7×10⁶ Nm/rad or 1.4×10⁷ Nm/rad maytherefore be of particular importance in some embodiments.

Relatively increasing the torsional stiffness of the carrier outside ofthe specified relationship may provide diminishing returns, or indeednegatively affect performance, due to the unnecessary size and/or weightof the stiffer carrier negating performance gains from reducing wind-up.

Each pin may have a soft connection to the carrier. The soft connectionmay be provided by one or more of a portion of the pin, a portion of thecarrier, and/or a separate component. The soft connection may be classedas a part of the pin for the purposes of assessing pin stiffness.

According to a twenty-fourth aspect, there is provided a method ofoperation of a gas turbine engine for an aircraft, the engine comprisingan engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; a fan located upstream of theengine core, the fan comprising a plurality of fan blades; and a gearboxarranged to receive an input from the core shaft and to output drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft. The gearbox is an epicyclic gearbox comprising a sun gear, aplurality of planet gears, a ring gear, and a planet carrier on whichthe planet gears are mounted.

A first carrier to pin stiffness ratio of:

$\frac{{{the}\mspace{14mu}{effective}\mspace{14mu}{linear}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}\mspace{14mu}}{{the}\mspace{14mu}{radial}\mspace{14mu}{bending}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}}$

-   -   is greater than or equal to 10, and optionally greater than or        equal to 15 (and/or optionally less than or equal to 40); or        a second carrier to pin stiffness ratio of:

$\frac{{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}\mspace{14mu}}{{the}\mspace{14mu}{tilt}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}}$

-   -   is greater than or equal to 24 (and optionally less than or        equal to 180).

The method comprises operating the gas turbine engine to providepropulsion under cruise conditions. The engine may be as described inthe twenty-second or twenty-third aspects.

The method may comprise driving the gearbox with an input torque of:

-   -   (i) greater than or equal to 10,000 Nm, and optionally of 10,000        to 50,000 Nm at cruise; and/or    -   (ii) greater than or equal to 28,000 Nm, and optionally of        28,000 to 135,000 Nm at Maximum Take-Off.

According to a twenty-fifth aspect, there is provided a propulsor for anaircraft, the propulsor comprising a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is an epicyclic gearbox arranged to receive an input from acore shaft driven by the power unit and to output drive to the fan so asto drive the fan at a lower rotational speed than the core shaft, andcomprises a sun gear, a plurality of planet gears, a ring gear, and aplanet carrier on which the planet gears are mounted. A first carrier topin stiffness ratio of:

$\frac{{{the}\mspace{14mu}{effective}\mspace{14mu}{linear}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}\mspace{14mu}}{{the}\mspace{14mu}{radial}\mspace{14mu}{bending}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}}$

-   -   is greater than or equal to 10, and optionally less than or        equal to 40.

According to a twenty-sixth aspect, there is provided a propulsor for anaircraft, the propulsor comprising a fan comprising a plurality of fanblades; a gearbox; and a power unit for driving the fan via the gearbox.The gearbox is an epicyclic gearbox arranged to receive an input from acore shaft driven by the power unit and to output drive to the fan so asto drive the fan at a lower rotational speed than the core shaft, andcomprises a sun gear, a plurality of planet gears, a ring gear, and aplanet carrier on which the planet gears are mounted. A second carrierto pin stiffness ratio of:

$\frac{{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}\mspace{14mu}}{{the}\mspace{14mu}{tilt}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}}$

-   -   is greater than or equal to 24, and optionally less than or        equal to 180.

The propulsor may have some or all of the features described above withrespect to the gas turbine engine, and may be a gas turbine engine insome embodiments.

In various other aspects of the invention, the specified boundaries onthe first and second carrier to pin stiffness ratios may be replaced by,or provided in addition to, specified boundaries on the product of thecomponents of the respective stiffness ratio, i.e. boundaries on thetorsional stiffness of the planet carrier multiplied by the tiltstiffness of a pin, or on the effective linear torsional stiffness ofthe planet carrier multiplied by the radial bending stiffness of a pin.The value of this product for the first carrier to pin stiffness ratio(i.e. the effective linear torsional stiffness of the planet carriermultiplied by the pin radial bending stiffness), in various aspects, maybe greater than or equal to 2.1×10¹⁸ N² m⁻², and optionally less than3.6×10²⁰ N²m⁻². Optionally, the value may be greater than or equal to5.8×10¹⁸ N²m², and optionally less than 1.7×10²⁰ N² m⁻². The value ofthis product for the second carrier to pin stiffness ratio 9 i.e. thetorsional stiffness of the planet carrier multiplied by the pin tiltstiffness), in various aspects, may be greater than or equal to 1.0×10¹⁵N² m²rad⁻², and optionally less than 4.7×10¹⁷ N²m²rad⁻².

Optionally, the value may be greater than or equal to 2.5×10¹⁵ N²m²rad⁻², and optionally less than 2.0×10¹⁷ N²m²rad⁻².

The skilled person would appreciate that gas turbine engine, method andpropulsor aspects may be formulated accordingly. The optional featuresof the aspects for the corresponding ratio may also apply to theseaspects.

In any of the preceding aspects, any one or more of the following mayapply as applicable:

The turbine may be a first turbine, the compressor a first compressor,and the core shaft a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

The planet carrier may comprise a forward plate and a rearward plate andpins extending therebetween. Each pin may be arranged to have a planetgear mounted thereon. The planet carrier may further comprise lugsextending between the forward and rearward plates, the lugs beingarranged to pass between adjacent planet gears.

The gearbox may comprise an odd number of planet gears, and optionallymay comprise 3, 5 or 7 planet gears.

The fan may have a fan diameter greater than 240 cm and less than orequal to 380 cm, and optionally greater than 300 cm and less than orequal to 380 cm.

The gearbox input shaft may provide a soft mounting for the sun gearsuch that some movement of the sun gear is facilitated. The core shaftmay comprise a more stiff section and a less stiff section, the lessstiff section providing the gearbox input shaft and being arranged tolie between the more stiff section and the sun gear and being arrangedto provide, or to contribute to, the soft mounting of the sun gear.

A gear ratio of the gearbox may be in any range disclosed herein, forexample in the range from 3.2 to 4.5, and optionally from 3.3 to 4.0.

A specific thrust of the engine at cruise may be in the range from 70 to90 NKg⁻¹s.

A bypass ratio at cruise may be in the range from 12.5 to 18; andoptionally from 13 to 16.

For any parameter or ratio of parameters X claimed or disclosed herein,a limit on the values that X can take that is expressed as “X is greaterthan or equal to Y” can alternatively be expressed as “1/X is less thanor equal to 1/Y”. Any of the ratios or parameters defined in the aspectsand statements above may therefore be expressed as “1/X is less than orequal to 1/Y” rather than “X is greater than or equal to Y”. Zero may betaken as the lower bound on the value of 1/X.

Various parameters of the gearbox, and/or of the engine more generally,may be adjusted to allow the engine to meet the specifications of thevarious aspects summarised above. Comments on various such parametersare provided below, with further examples of ways in which these may beadjusted provided later in the description of the components.

One or more of gearbox size, gearbox geometry (including presence orabsence of lugs in the carrier, and the number, size, and/or shape ofany lugs present), and material choice, amongst other factors, may beselected or adjusted to achieve a desired carrier stiffness. Thematerials of which the carrier is made (often steels) may, for example,have a Young's modulus in the range from 100 to 250 GPa, or 105 to 215GPa, and optionally around 210 GPa—different grades of steel, or othertypes of metal, may be selected to achieve different stiffnesses for thesame size and geometry. For example, steels with a Young's modulus inthe range 190 to 215 GPa, titanium alloys with a Young's modulus in therange 105 to 120 GPa, or a metal such as titanium with a Young's modulusof around 110 GPa may be used in various embodiments.

Flexibility of the carrier (effectively the inverse of stiffness) allowschanges in alignment of the gears and bearings—the inventor appreciatedthat a certain amount of flexibility in some places may advantageouslyallow manufacturing misalignments to be corrected in use, that a certainmisalignment may be tolerated, and that a larger misalignment coulddeleteriously affect running of the engine, and discovered variousstiffness relationships to capture the advantages of appropriatestiffness ranges.

One or more of material choice, pin geometry (e.g. diameter), pinmounting design, and internal pin structure (e.g. solid or hollow) maybe selected or adjusted to achieve a desired pin stiffness. Pinmaterials may often be steels (often with a Young's modulus of 100 to250 GPa, and optionally around 210 GPa) and one or more different steelgrades may be selected to adjust stiffness.

Some flexibility of the pins may be provided to allow correction ofplanet misalignment, but too much flexibility may create damagingmisalignments. Increasing pin stiffness too far may result in excessivesize and/or weight reducing overall performance.

Turning to the gearbox input shaft, the inventor has discovered that thetorsional stiffness of the gearbox input shaft has an effect on thetorsional stiffness of the whole transmission, but a relatively minimaleffect on gearbox operation as torsional deflection results in wind uponly, and no misalignment of gears. The gearbox input shaft maytherefore have a lower torsional stiffness than the carrier withoutdeleterious effects.

Similar considerations may apply to the fan shaft (the gearbox outputshaft).

The inventor realised that decreasing the torsional stiffnesses of theshafts below the ranges defined herein may result in deleterious torsionvibrations at low modal frequencies (the skilled person would appreciatethat the lower modal frequency whirl modes have largeramplitudes/deflections than the higher modes, and so are more importantto avoid), whilst increasing the torsional stiffness above the rangesdefined herein may result in excessive size and/or weight of the shaftwithout a corresponding improvement in performance. One or more of shaftdiameter, material(s), and wall thickness may be adjusted so as toobtain shaft stiffnesses in the desired ranges.

Turning to gearbox size, and in particular to ring gear pitch circlediameter (PCD) as a measure of gearbox size, the inventor appreciatedthat an optimal PCD may be selected by considering the relationshipbetween improved performance due to improved use of the lever effect forlarger gearbox sizes, and the effect of increased drag for largergearbox sizes (diminishing returns on the improved lever effect from thelarger size above a certain PCD, and increased size and weight of thelarger size). Ring gear materials may be selected to ensure that amaximum expected torque density for the PCD size would be well withintolerance limits.

One or more of gearbox support material(s) and geometry may be adjustedto achieve a desired torsional stiffness. The inventor appreciated thatgearbox support torsional stiffness may be selected to be high enough tosuppress torque ripple effects, so maintaining gearbox movements withinacceptable bounds, whilst avoiding the addition of unnecessary sizeand/or weight.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

The gas turbine engine may comprise a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft. The input to thegearbox may be directly from the core shaft, or indirectly from the coreshaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. Thefan tip loading may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20.

The bypass ratio may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of form 12 to 16, 13 to 15, or 13 to14. The bypass duct may be substantially annular. The bypass duct may beradially outside the core engine. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

As used herein, a maximum take-off (MTO) condition has the conventionalmeaning. Maximum take-off conditions may be defined as operating theengine at International Standard Atmosphere (ISA) sea level pressure andtemperature conditions+15° C. at maximum take-off thrust at end ofrunway, which is typically defined at an aircraft speed of around 0.25Mn, or between around 0.24 and 0.27 Mn. Maximum take-off conditions forthe engine may therefore be defined as operating the engine at a maximumtake-off thrust (for example maximum throttle) for the engine atInternational Standard Atmosphere (ISA) sea level pressure andtemperature +15° C. with a fan inlet velocity of 0.25 Mn.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000m to15000m, for example in the range of from 10000m to 12000m, for examplein the range of from 10400m to 11600m (around 38000 ft), for example inthe range of from 10500m to 11500m, for example in the range of from10600m to 11400m, for example in the range of from 10700m (around 35000ft) to 11300m, for example in the range of from 10800m to 11200m, forexample in the range of from 10900m to 11100m, for example on the orderof 11000m. The cruise conditions may correspond to standard atmosphericconditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number (Mn) of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

Whilst in the arrangements described herein the source of drive for thepropulsive fan is provided by a gas turbine engine, the skilled personwill appreciate the applicability of the gearbox configurationsdisclosed herein to other forms of aircraft propulsor comprisingalternative drive types. For example, the above-mentioned gearboxarrangements may be utilised in aircraft propulsors comprising apropulsive fan driven by an electric motor. In such circumstances, theelectric motor may be configured to operate at higher rotational speedsand thus may have a lower rotor diameter and may be more power-dense.The gearbox configurations of the aforesaid aspects may be employed toreduce the rotational input speed for the fan or propeller to allow itto operate in a more favourable efficiency regime. Thus, according to anaspect, there is provided an electric propulsion unit for an aircraft,comprising an electric machine configured to drive a propulsive fan viaa gearbox, the gearbox and/or its inputs/outputs/supports being asdescribed and/or claimed herein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect.

Furthermore, except where mutually exclusive, any feature or parameterdescribed herein may be applied to any aspect and/or combined with anyother feature or parameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a schematic diagram illustrating automatic load-shareadjustments within a gearbox;

FIG. 5 is a schematic diagram illustrating radial bending stiffness of acantilevered beam;

FIG. 6 is a schematic diagram illustrating tilt stiffness of acantilevered beam;

FIG. 7 is a schematic diagram illustrating torsional stiffness of acantilevered beam;

FIG. 8 is a schematic diagram illustrating radial bending stiffness of acantilevered beam, with its moveable end slidably mounted on a plane;

FIG. 9 is a schematic diagram illustrating radial bending stiffness ofthe carrier;

FIG. 10 is a schematic diagram illustrating tilt stiffness of thecarrier, and more specifically the determination of an effective lineartilt stiffness for the carrier;

FIG. 11 is a schematic diagram illustrating tilt stiffness of thecarrier;

FIG. 12 is a schematic diagram illustrating torsional stiffness of thecarrier in side view;

FIG. 13 is a schematic diagram illustrating torsional stiffness of analternative carrier in front view;

FIG. 14 is a schematic diagram illustrating torsional stiffness of thecarrier of FIG. 13;

FIG. 15 is a schematic diagram illustrating a front view of a carriercomprising lugs;

FIG. 16 is a schematic diagram illustrating radial bending stiffness ofunjointed pins;

FIG. 17 is a schematic diagram illustrating radial bending stiffness ofjointed pins;

FIG. 18 is a schematic diagram illustrating tilt stiffness of pins;

FIG. 19 is a schematic diagram illustrating the core shaft, and inparticular the gearbox input shaft;

FIG. 20 is a schematic diagram illustrating the torsional stiffness ofthe gearbox input shaft;

FIG. 21 includes side and radial views of the gearbox supportillustrating torsional stiffness of the gearbox support;

FIG. 22 is a schematic diagram illustrating a portion of an engine witha star gearbox;

FIG. 23 is a schematic diagram illustrating connection of the fan shaftto a star gearbox;

FIG. 24 is a schematic diagram illustrating connection of the fan shaftto a planetary gearbox;

FIG. 25 is a schematic diagram illustrating fan shaft torsionalstiffness in an engine with a star gearbox;

FIG. 26 is a graph of displacement against load, illustrating an elasticregion within which stiffnesses of components may be determined; and

FIG. 27 illustrates a method.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

The linkages 36 may be referred to as a fan shaft 36, the fan shaft 36optionally comprising two or more shaft portions coupled together. Forexample, the fan shaft 36 may comprise a gearbox output shaft portion 36a extending from the gearbox 30 and a fan portion 36 b extending betweenthe gearbox output shaft portion and the fan 23. In the embodiment shownin FIGS. 1 and 2, the gearbox 30 is a planetary gearbox and the gearboxoutput shaft portion 36 a is connected to the planet carrier 34—it maytherefore be referred to as a carrier output shaft 36 a. In stargearboxes 30, the gearbox output shaft portion 36 a may be connected tothe ring gear 38—it may therefore be referred to as a ring output shaft36 a. In the embodiment shown in FIGS. 1 and 2, the fan portion 36 b ofthe fan shaft 36 connects the gearbox output shaft portion 36 a to thefan 23. The output of the gearbox 30 is therefore transferred to the fan23, to rotate the fan, via the fan shaft 36. In alternative embodiments,the fan shaft 36 may comprise a single component, or more than twocomponents. Unless otherwise indicated or apparent to the skilledperson, anything described with respect to an engine 10 with a stargearbox 30 may equally be applied to an engine with a planetary gearbox30, and vice versa.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular. In the describedarrangement, the carrier 34 comprises two plates 34 a, 34 b; inparticular a forward plate 34 a and a rearward plate 34 b (see, forexample, FIG. 9). Each plate 34 a, 34 b extends in a radial plane, withthe forward plate 34 a lying further forward in the engine 10/closer tothe fan 23 than the rearward plate 34 b.

The carrier 34 may take any suitable form. For example, the carrier mayor may not be symmetric about its axial mid-point. Purely by wayexample, in a described arrangement, the carrier 34 is not symmetricabout its axial mid-point, but rather the rearward plate 34 b is stifferthan the forward plate 34 a (for example by 50 to 300%) to compensatefor an asymmetric torque variation across the gearbox 30. In someembodiments, no forward plate 34 a may be provided, or only a smallerforward plate 34 a. In some embodiments, the plates 34 a, 34 b of thecarrier 34 may have equal stiffnesses (for example, in various planetarygearbox arrangements; stiffer rearward plates 34 b may be preferred insome star gearbox arrangements).

A plurality of pins 33 extend across the carrier 34 (between the forwardand rearward plates 34 a,b in the arrangement being described), asshown, for example, in FIGS. 9 to 18. The pins 33 form a part of thecarrier 34. Each pin 33 has a planet gear 32 mounted thereon. Asreferred to herein, reference to a planet gear 32 includes the gear 32mounted on the pin 33 regardless of whether the gearbox is a so-called“star” arrangement (such as that shown in FIG. 22), or a “planetarygearbox” (such as that shown in FIG. 2).

The carrier stiffness in a region at each of the front and rear ends ofeach pin 33 may be arranged to be relatively low in the embodimentsbeing described, to facilitate a more even load distribution; i.e. toimprove load-share factor. This may be described as soft mountings foreach pin 33. The soft mountings 33 a, 33 b may allow some movement ofthe pins 33 relative to each other, and relative to the carrier plates34 a, 34 b, so allowing differences between planet gears 34, or othermanufacturing defects, to be accommodated without a significantdifference in load between different planet gears 34.

Such soft mountings 34 a, 34 b may be provided by a portion of the pin33, by a separate component, and/or by a portion of the respectivecarrier plate 34 a, 34 b, in various embodiments. The soft mountings 34a, 34 b may be designed to accommodate movements to address one or moreof carrier bearing location accuracy and clearance, planet pin runout ofbearing surface to mounting feature(s), planet gear teeth to bearingrunout, planet gear teeth spacing and thickness variation/manufacturingtolerances, sun gear teeth spacing and thickness variation/manufacturingtolerances, and/or gearbox input shaft mainline bearing locationaccuracy and clearance, or the likes. For example, in variousembodiments the soft mountings 34 a, 34 b may be arranged to allowaround 500 μm of pin movement.

Pin size, design and/or material may be adjusted to provide appropriatestiffnesses to the carrier 34.

In some arrangements, such as that shown in FIG. 15, lugs 34 c may beprovided, extending between the carrier plates 34 a, 34 b and past theplanet gears 32. The presence/absence of lugs 34 c, and the number,shape, and/or material(s) of the lug(s) may vary in various embodiments,and may be adjusted to provide appropriate stiffnesses to the carrier34.

Use of flexibility within the gearbox 30 to improve load-share isillustrated schematically in FIG. 4, which shows a planetary gearbox 30with three planet gears 32 a, 32 b, 32 c (misalignments are exaggeratedin this Figure for clarity of explanation). In this example, the sungear 28 is slightly off-centre with respect to the ring gear 38, and inparticular is closer to two planet gears 32 a and 32 b than it is to thethird planet gear 32 c. In the schematic example shown, there is nocontact between the third planet gear 32 c and the sun gear 28, leavingthe other two planet gears 32 a, 32 b to take 50% of the load each,rather than one third each as would be expected for an even loaddistribution. This relatively extreme example is provided for ease ofreference only—in reality, situations in which contact with one planetgear 32 c is reduced, but not completely eliminated, would be morelikely, for example leading to a percentage load-share of 20:40:40,26:37:37, or 31:34:34 or the likes rather than the ideal even load sharefraction of ⅓: ⅓: ⅓ (i.e. 33:33:33, as a percentage load share, roundedto the nearest integer).

In the example shown in FIG. 4, each of the two planet gears 32 a, 32 bin contact with the sun gear 28 exerts a force F_(a), F_(b) on the sungear 28. The resultant force, F_(R), on the sun gear 28 pushes the sungear 28 towards the third planet gear 32 c, so re-establishing contactand making the load-share between planets 32 more even. A soft mountingof the sun gear 28/flexibility in the core input shaft 26 facilitatesthis re-balancing. Such a soft mounting of the sun gear 28 may bedesigned to accommodate movements to address one or more of carrierbearing location accuracy and clearance, planet and/or sun gear teethspacing and thickness variation/manufacturing tolerances, and/or gearboxinput shaft mainline bearing location accuracy and clearance, or thelikes. For example, in various embodiments such soft mountings may bearranged to allow around 1000 μm of sun gear movement.

The skilled person would appreciate that a similar effect would apply ifone of the planet gears 32 were closer to the sun gear 28 than theothers; pushing the relevant planet gear 32 back towards the ring gear38, or if one of the planet gears 32 were larger or smaller than theothers. Soft mounting of the pins 33/flexibility in the carrier 34facilitates this re-balancing. Small variations between planet gears 32and/or misalignments of pins 33 or shafts 26 may therefore beaccommodated by flexibility within the gearbox 30. Having an odd numberof planet gears 32 (e.g. 3, 5 or 7 planet gears) may facilitate thisautomatic re-distribution of load-share, although even numbers of planetgears may be used in some arrangements.

The following general definitions of stiffnesses may be used herein:

Radial Bending Stiffness

A radial bending stiffness is a measure of deformation caused by a givenforce applied in any one selected radial direction (i.e. any directionperpendicular to and passing through the engine axis). The radialbending stiffness is defined with reference to FIG. 5 in terms of thedeformation of a cantilevered beam 401. As illustrated in FIG. 5, aforce, F, applied at the free end of the beam in a directionperpendicular to the longitudinal axis of the beam causes a linearperpendicular deformation, δ. The radial bending stiffness is the forceapplied for a given linear deformation i.e. F/δ. In the presentapplication, the radial bending stiffness is taken relative to therotational axis of the engine 9, and so relates to the resistance tolinear deformation in a radial direction of the engine caused by aradial force. The beam, or equivalent cantilevered component, extendsalong the axis of rotation of the engine, the force, F, is appliedperpendicular to the axis of the engine, along any radial direction, andthe displacement 6 is measured perpendicular to the axis of rotation,along the line of action of the force. The radial bending stiffness asdefined herein has SI units of N/m. In the present application, unlessotherwise stated, the radial bending stiffness is taken to be afree-body stiffness i.e. stiffness measured for a component in isolationin a cantilever configuration, without other components present whichmay affect its stiffness. When the force is applied perpendicular to thecantilevered beam, and at the free end of the beam, the resultantcurvature is not constant but rather increases towards the fixed end ofthe beam.

For some components, the beam may more appropriately be constrained tomove in a particular way as is described in more detail for the specificexample of pins 33 below.

Tilt Stiffness

A tilt stiffness is defined with reference to FIG. 6, which shows theresulting deformation of a cantilevered beam 401 under a moment, M,applied at its free end. The tilt stiffness is a measure of theresistance to rotation of a point on the component at which a moment isapplied. As can be seen in FIG. 6, an applied moment at the free end ofthe cantilevered beam causes a constant curvature along the length ofthe beam between its free and fixed ends. The applied moment, M, causesa rotation θ of the point at which it is applied. The tilt stiffness asdefined herein therefore has SI units of Nm/radian.

The tilt stiffness may be expressed as an effective linear tiltstiffness for a component having a given radius by expressing the tiltstiffness in terms of a pair of equal and opposite forces, F, acting ateither end of that radius (rather than the moment) and the arcdisplacement at that radius (i.e. displacement measured along acircumference of a circle having that radius). An approximate or overalltilt angle, a, may be defined for the purposes of calculating theeffective linear stiffness. The arc displacement may be referred to asrα. The effective linear tilt stiffness is given by the ratio ofeffective force divided by the displacement, F/rα and has the units N/m.

Torsional Stiffness

FIG. 7 illustrates the definition of the torsional stiffness of a shaft401 or other body. A torque, τ, applied to the free end of the beamcauses a rotational deformation, θ (e.g. twist) along the length of thebeam. The torsional stiffness is the torque applied for a given angle oftwist i.e. τ/θ. The torsional stiffness has SI units of Nm/rad.

An effective linear torsional stiffness may be determined for acomponent having a given radius. The effective linear torsionalstiffness is defined in terms of an equivalent tangential force appliedat a point on that radius (with magnitude of torque divided by theradius) and the distance δ (with magnitude of the radius multiplied byθ) moved by a point corresponding to the rotational deformation θ of thecomponent.

More specific definitions of stiffnesses relating to embodimentsdescribed herein are provided below for ease of understanding.

Carrier Stiffnesses

The planet carrier 34 holds the planet gears 32 in place. In variousarrangements including the embodiment being described, the planetcarrier 34 comprises a forward plate 34 a and a rearward plate 34 b, andpins 33 extending between the plates, as illustrated in FIGS. 9 to 17.The pins 33 are arranged to be parallel to the engine axis 9. Inalternative embodiments, a plate 34 b may be provided on only oneside—no plate or only a partial plate may be provided on the other sideof the carrier 34. In the embodiment shown in FIG. 15, the carrier 34additionally comprises lugs 34 c, which may also be referred to aswedges or a web, between the forward and rearward plates 34 a, 34 b. Thelugs 34 c may increase the overall stiffness of the carrier 34.

The stiffness of the carrier 34 is selected to be relatively high toreact centrifugal forces and/or to maintain gear alignment. The skilledperson would appreciate that stiffness is a measure of the displacementthat results from any applied forces or moments, and may not relate tostrength of the component. Hence to react a high load, any stiffness isacceptable so long as the resulting displacement is tolerable. How higha stiffness is desired to keep a displacement within acceptable limitstherefore depends on position and orientation of the gears, which isgenerally referred to as gear alignment (or mis-alignment).

Carrier Radial Bending Stiffness

In the embodiment being described, the carrier radial bending stiffnessis determined treating the carrier 34 as a free-body fixedly mounted atone plate 34 b, with a (radial) force, F, applied at the axial positionof the axial centre point of the other plate 34 a. This is illustratedin FIG. 9, with arrow F indicating the (radial) force on the plate 34 aand δ illustrating the (radial) displacement of the plate 34 a. Theforce, F, is shown acting along a line that passes through the engineaxis 9. In embodiments with only one plate 34 a, the ends of the pins 33further from the single plate 34 a are held in place instead.

In various embodiments, a radial bending stiffness of the carrier 34 maybe equal to or greater than 1.20×10⁹ N/m, and optionally equal to orgreater than 2.0×10⁹ N/m. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the radial bending stiffness of the carrier 34 may be equal to orgreater than 1.5×10⁹ N/m, and optionally may be equal to or greater than2×10⁹ N/m (and optionally may be equal to 2.30×10⁹ or 3.85×10⁹ N/m). Insome embodiments, for example in embodiments in which the fan diameteris in the range from 330 to 380 cm, the radial bending stiffness of thecarrier 34 may be equal to or greater than 2.0×10⁹ N/m and optionallymay be equal to or greater than 3×10⁹ N/m (and optionally may be equalto 3.92×10⁹ N/m or 7.70×10⁹ N/m).

In various embodiments, the radial bending stiffness of the carrier 34is in the range from 1.20×10⁹ to 1×10¹² N/m, and optionally in the rangefrom 2.0×10⁹ to 1.5×10¹¹ N/m. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the radial bending stiffness of the carrier 34 may be in the rangefrom 1.5×10⁹ to 5×10¹⁰ N/m, and optionally may be in the range from2×10⁹ to 5×10⁹ N/m or from 1.9×10⁹ to 2.7×10⁹ N/m (and optionally may beequal to 3.85×10⁹ N/m or to 2.30×10⁹ N/m). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the radial bending stiffness of the carrier 34 may be inthe range from 2.0×10⁹ to 1.6×10¹¹ N/m and optionally may be in therange from 3.0×10⁹ to 9.0×10⁹ N/m (and optionally may be equal to7.70×10⁹ N/m or to 3.92×10⁹ N/m).

The carrier 34 serves to locate the planet gears 32 within the gearbox30 and to reduce or avoid misalignment. The skilled person wouldappreciate that the desired carrier stiffness may be obtained in variousdifferent ways, for example by adjusting one or more of carriermaterial(s) and carrier geometry as appropriate. For a given material,the stiffness may be a function of carrier and gearbox size, and gearboxconfiguration, for example. Planet gear number and gearbox ratio mayalso be adjusted to arrive at the desired stiffness. For example, thegear ratio of the gearbox 30 may be changed, so causing the planet gearspacing, and in some cases the number of planet gears 32, to change. Thechange in planet gear spacing may provide more (or less) space for lugs34 c between the gears 32, and the sizes and shapes of these lugs 34 cmay be adjusted to achieve the desired carrier stiffness.

Carrier Tilt Stiffness

Carrier tilt stiffness is a measure of the resistance of the carrier 34to an applied moment, M, as illustrated in FIG. 10. The axis of themoment is perpendicular to the engine axis 9. Two points of the carrier34 are selected for measuring tilt stiffness: a forward point at theaxial position of the axial centre point of the forward plate 34 a andrearward point at the axial position of the axial centre point of therearward plate 34 b. The rearward plate 34 b is held to be rigid andnon-rotating, as illustrated by the diagonal lines in FIG. 10.

In response to the applied moment, M, which is an anticlockwise momentin the example shown, but could be a clockwise moment in other examples,the carrier 34 bends through an angle θ—the angle θ is not constant ateach point along the length of the carrier 34 as the carrier does nothave a constant section. θ may therefore be measured between a lineparallel to the engine axis 9 and passing through the rearward point andthe forward point pre-deformation (perpendicular to forward and rearfaces of the carrier plates) and a line passing through the forwardpoint and perpendicular to forward and rear faces of the forward carrierplate (no longer parallel to the engine axis) after the deformation.This is shown in FIG. 11.

The carrier 34 bends through an overall angle α, resulting in an arcdisplacement 6. The angle, a, is measured between a line parallel to theengine axis 9 and passing through the rearward point and the forwardpoint pre-deformation (as for θ) and a line passing through the forwardpoint and the rearward point after the deformation (in contrast to θ).The values of θ and a may therefore be different.

An effective linear tilt stiffness can therefore be defined for thecarrier 34 as described above. The radius, r, chosen for the definitionof the effective linear tilt stiffness is that of a circle centred on apoint on the surface of the plate 34 b held to be rigid and passingthrough an original axial centrepoint of the forward plate 34 a and thesame point post-deformation, as shown in the close-up portion of FIG.10. The two equal and opposite forces, F, are marked on the schematicforce diagram shown on the bottom right of FIG. 10, one at the centre ofthe circle of radius r, and one at the far end of that radius—themagnitude of F is selected based on the moment applied. The overallangle, a, is measured between a first radius before the deformation anda second radius after the deformation. The axial centrepoint at theradial position of the lower edge of the forward plate 34 a is selectedfor ease of reference—any other point along the axial centreline of theforward plate 34 a may be selected equivalently (e.g. the point ofapplication of the moment). The arc displacement (arc distance) δ isequal to rα.

In various embodiments, the tilt stiffness of the carrier 34 is in therange is greater than or equal to 6.00×10⁸ Nm/radian, and optionallygreater than or equal to 1.3×10⁹ Nm/radian. In some embodiments, forexample in embodiments in which the fan diameter is in the range from240 to 280 cm, the tilt stiffness of the carrier 34 may be greater thanor equal to 2.2×10⁹ Nm/radian, and optionally may greater than or equalto 2.4×10⁹ Nm/radian (and optionally may be equal to 2.71×10⁹Nm/radian). In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the tilt stiffness ofthe carrier 34 may be greater than or equal to 2.3×10⁹ Nm/radian andoptionally may be greater than or equal to 3.0×10⁹ Nm/radian (andoptionally may be equal to 5.70×10⁹ Nm/radian).

In various embodiments, the tilt stiffness of the carrier 34 is in therange from 6.00×10⁸ to 2.80×10¹¹ Nm/radian, and optionally in the rangefrom 1.3×10⁹ to 1.2×10¹¹ Nm/radian. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the tilt stiffness of the carrier 34 may be in the range from2.2×10⁹ to 1.4×10¹¹ Nm/radian, and optionally may be in the range from2.4×10⁹ to 5.0×10⁹ Nm/radian (and optionally may be equal to 2.71×10⁹Nm/radian). In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the tilt stiffness ofthe carrier 34 may be in the range from 2.3×10⁹ to 2.8×10¹¹ Nm/radianand optionally may be in the range from 3.0×10⁹ to 9.0×10⁹ Nm/radian(and optionally may be equal to 5.70×10⁹ Nm/radian).

In various embodiments, the effective linear tilt stiffness of thecarrier 34 is greater than or equal to 3.40×10⁹ N/m, and optionallygreater than or equal to 8.0×10⁹ N/m. In some embodiments, for examplein embodiments in which the fan diameter is in the range from 240 to 280cm, the effective linear tilt stiffness of the carrier 34 may be greaterthan or equal to 1.4×10¹⁰ N/m, and optionally may be greater than orequal to 1.42×10¹⁰ N/m (and optionally may be equal to 1.68×10¹⁰ N/m).In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the effective linear tiltstiffness of the carrier 34 may be greater than or equal to 1.5×10¹⁰ N/mand optionally may be greater than or equal to 3.0×10¹⁰ N/m, andoptionally greater than or equal to 7.0×10¹⁰ N/m (and optionally may beequal to 8.36×10¹⁰ N/m).

In various embodiments, the effective linear tilt stiffness of thecarrier 34 is in the range from 3.40×10⁹ to 4.20×10¹² N/m, andoptionally in the range from 8.0×10⁹ to 1.7×10¹² N/m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the effective linear tilt stiffness of thecarrier 34 may be in the range from 1.4×10¹⁰ to 8.4×10¹¹ N/m, andoptionally may be in the range from 1.42×10¹⁰ to 2.72×10¹⁰ N/m (andoptionally may be equal to 1.68×10¹⁰ N/m). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the effective linear tilt stiffness of the carrier 34 maybe in the range from 1.5×10¹⁰ to 4.2×10¹² N/m and optionally may be inthe range from 3.0×10¹⁰ to 1.0×10¹¹ N/m, and optionally from 7.0×10¹⁰ to1.0×10¹¹ N/m (and optionally may be equal to 8.36×10¹⁰ N/m).

As discussed above, excess tilt deflections of the carrier 34 may bemore damaging than radial bending or torsional deflections for the samemagnitude of deflection as tilt deflections may produce two compoundingeffects—firstly, load share may worsen, with some planet gears 32 takinga larger share of the load than others, and secondly face distributionof that load shifts. The larger force on a particular planet gear 32 istherefore concentrated on one side of the gear rather than equallydistributed across the tooth. The increased load on that gear 32 and theincreased concentration of that load may therefore damage the gearteeth.

Carrier Torsional Stiffness

Carrier torsional stiffness is a measure of the resistance of thecarrier 34 to an applied torque, τ, as illustrated in FIG. 12 (axialcross-section) and FIGS. 13 to 15 (radial cross-section). The axis ofthe torque is parallel to the engine axis 9. The diagonal lining of theplate 34 b at the rearward end of the carrier 30 indicates that plate 34b is treated as rigid and non-rotating (as for a cantilever beammounting). In embodiments with only one plate 34 a, the ends of the pins33 (and of the lugs 34 c if present) further from the single plate 34 aare held in place instead.

The torque, τ, is applied to the carrier 34 (at the position of theaxial mid-point of the forward plate 34 a) and causes a rotationaldeformation, θ (e.g. twist) along the length of the carrier 34. Thetwist causes the carrier 34 to “wind up” as the ends of the pins 33 (andof the lugs 34 c if present) are held at a fixed radius on the carrierplates 34 a, 34 b.

The angle through which a point on an imaginary circle 902 on theforward plate 34 a passing through the longitudinal axis of each pin 33moves is θ, where θ is the angle measured in radians. The imaginarycircle 902 may be referred to as the pin pitch circle diameter (pinPCD). The pin PCD may be in the range from 0.38 to 0.65 m, for examplebeing equal to 0.4 m or 0.55 m. An effective linear torsional stiffnesscan therefore be defined for the carrier 34 as described above, usingthe radius r of the imaginary circle 902 (e.g. as illustrated in FIG.13).

In various embodiments, the torsional stiffness of the carrier 34 isgreater than or equal to 1.60×10⁸ Nm/rad, and optionally greater than orequal to 2.7×10⁸ Nm/rad. In some embodiments, for example in embodimentsin which the fan diameter is in the range from 240 to 280 cm, thetorsional stiffness of the carrier 34 may be greater than or equal to1.8×10⁸ Nm/rad, and optionally may be greater than or equal to 2.5×10⁸Nm/rad (and optionally may be equal to 4.83×10⁸ Nm/rad). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the torsional stiffness of the carrier 34may greater than or equal to 6.0×10⁸ Nm/rad and optionally may begreater than or equal to 1.1×10⁹ Nm/rad (and optionally may be equal to2.17×10⁹ Nm/rad).

In various embodiments, the torsional stiffness of the carrier 34 is inthe range from 1.60×10⁸ to 1.00×10¹¹ Nm/rad, and optionally in the rangefrom 2.7×10⁸ to 1×10¹⁰ Nm/rad. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the torsional stiffness of the carrier 34 may be in the range from1.8×10⁸ to 4.8×10⁹ Nm/rad, and optionally may be in the range from2.5×10⁸ to 6.5×10⁸ Nm/rad (and optionally may be equal to 4.83×10⁸Nm/rad). In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the torsional stiffnessof the carrier 34 may be in the range from 6.0×10⁸ to 2.2×10¹⁰ Nm/radand optionally may be in the range from 1.1×10⁹ to 3.0×10⁹ Nm/rad (andoptionally may be equal to 2.17×10⁹ Nm/rad).

In various embodiments, the effective linear torsional stiffness of thecarrier 34 may be greater than or equal to 7.00×10⁹ N/m, and optionallygreater than or equal to 9.1×10⁹ N/m. In some embodiments, for examplein embodiments in which the fan diameter is in the range from 240 to 280cm, the effective linear torsional stiffness of the carrier 34 may begreater than or equal to 7.70×10⁹ N/m. In other such embodiments, theeffective linear torsional stiffness of the carrier 34 may be greaterthan or equal to 9.1×10⁹ N/m, optionally greater than or equal to1.1×10¹⁰ N/m (and optionally may be equal to 1.26×10¹⁰ N/m). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the effective linear torsional stiffnessof the carrier 34 may be greater than or equal to 1.2×10¹⁰ N/m andoptionally may be greater than or equal to 2.1×10¹⁰ N/m (and optionallymay be equal to 2.88×10¹⁰ N/m).

In various embodiments, the effective linear torsional stiffness of thecarrier 34 may be in the range from 7.00×10⁹ to 1.20×10¹¹ N/m, andoptionally in the range from 9.1×10⁹ to 8.0×10¹⁰ N/m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the effective linear torsional stiffnessof the carrier 34 may be in the range from 9.1×10⁹ to 6.0×10¹⁰ N/m, andoptionally may be in the range from 7×10⁹ to 2×10¹⁰ N/m, or from 8.5×10⁹to 2×10¹⁰ N/m (and optionally may be equal to 1.26×10¹⁰ N/m). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the effective linear torsional stiffnessof the carrier 34 may be in the range from 1.2×10¹⁰ to 1.2×10¹¹ N/m andoptionally may be in the range from 1.0×10¹⁰ to 5.0×10¹⁰ N/m (andoptionally may be equal to 2.88×10¹⁰ N/m).

The torsional stiffness of the carrier 34 may be controlled so as to bewithin a desired range by adjusting one or more parameters, includingcarrier material(s), carrier geometry, and the presence or absence oflugs.

Planet Pin Stiffnesses

The planet carrier 34 holds the planet gears 32 in place. The pins 33serve to locate each planet gear 32 on the carrier 34 and to reduce oravoid misalignment of the planet gears 32. In the embodiment shown inFIG. 16, five planet gears 32 are provided in the gearbox 30, with onepin 33 for each planet gear 32. In alternative embodiments, a differentnumber of planet gears 32 may be provided, for example 3, 4, 5, 6, 7, 8or 9 planet gears 32. The skilled person would appreciate that an oddnumber of planet gears (e.g. 3, 5, 7, 9) may improve load and/or stresssharing within the gearbox 30. In particular, the skilled person wouldappreciate that the use of an odd number of planet gears may improve thedynamic loading in gearboxes which have sun gear tooth numbers which arenot whole numbers divisible by the number of planets (e.g. 41 sun gearteeth, 5 planets) as in the embodiment being described. In alternativeembodiments in which the sun gear tooth number is wholly divisible bythe number of planets (e.g. sun gear 40 teeth, 5 planets) there may beless or no dynamic benefit.

Each planet gear 32 is mounted to the rest of the planet carrier 34 by apin 33, also referred to as a planet pin. The planet pins 33 are mountedon the carrier 34 such that they move with the carrier 34 as the carrierrotates in a planetary arrangement, or remain in place with anon-rotating carrier 34 in a star arrangement. The pins 33 may bereferred to as axles/supports for the planets 32. A lower stiffness ofthe planet pins 33 may reduce differential load and improve load sharefactor relative to a higher stiffness of the planet pins 33.

The pins 33 may be connected to the carrier plates 342, 34 b in anydesired manner. For example, in the embodiment being described, the pins33 are provided with a soft connection 31 a, 31 b to each carrier plate34 a, 34 b. The connection is described as soft as it is arranged tofacilitate some movement of the pins 33, which may help to improve loadshare. Such a soft connection may be formed by the plate 34 a,bitself—for example having cut-away portions of material to provide somemovement of the plate—or by a part of the pin, or by a separatecomponent. The soft connections 31 a, 31 b are classed as a part of thepin 33 for the assessment of the stiffnesses described herein. The softconnections 33 a, 33 b are illustrated only for the pin of interest inFIGS. 16 and 17, but the skilled person would appreciate that each pin33 would have an equivalent connection.

Planet Pin Radial Bending Stiffness

The radial bending stiffness of a pin 33 may be measured in differentways, for example depending on the pin design. The soft connections 31a, 31 b of the pin 33, if present, are classed as a part of the pin 33.For an unjointed pin 33, as shown in FIG. 16, each pin 33 may be treatedas being rigidly mounted on one carrier plate 34 b (illustrated by thediagonal lines on the carrier plate 34 b). The other end of each pin maybe treated as being slidably mounted on the other carrier plate 34 a,such that the end of the pin 33 can slide along the plate 34 a, butcannot move away from the plate 34 a. A radial force, F, is then appliedat the position on the pin corresponding to the sliding plane, and aresultant radial displacement, δ, at that axial position is measured.

For a pin 33 with two separate shaft portions 33 a, 33 b and a joint 33c (e.g. a ball joint) between the two, as shown in FIG. 17, the pin 33may instead be treated as being rigidly mounted on both carrier plates34 a,34 b (illustrated by the diagonal lines on the carrier plates 34 a,34 b). A radial force, F, is then applied at the position on the pin 33corresponding to the axial centre point of the pin, and a resultantradial displacement, δ, at that axial position is measured. Inembodiments with a non-central joint 33 c, the axial centre point of thejoint 33 c may be selected in place of the axial centre point of the pin33.

In various embodiments, the radial bending stiffness of the pin 33 isgreater than or equal to 3.00×10⁸ N/m, and optionally greater than orequal to 6.3×10⁸ N/m. In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 240 to 280 cm, the radialbending stiffness of the pin 33 shaft may be greater than or equal to6.3×10⁸ N/m, and optionally may be greater than or equal to 6.7×10⁸ N/m(and optionally may be equal to 7.70×10⁸ N/m). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the radial bending stiffness of the pin 33 may be greaterthan or equal to 9.0×10⁸ N/m and optionally may be greater than or equalto 1.0×10⁹ N/m (and optionally may be equal to 1.54×10⁹ N/m).

In various embodiments, the radial bending stiffness of the pin 33 is inthe range from 3.00×10⁸ to 3.00×10⁹ N/m, and optionally in the rangefrom 6.3×10⁸ to 2.5×10⁹ N/m. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the radial bending stiffness of the pin 33 shaft may be in the rangefrom 6.3×10⁸ to 1.5×10⁹ N/m, and optionally may be in the range from6.7×10⁸ to 8.7×10⁸ N/m (and optionally may be equal to 7.70×10⁸ N/m). Insome embodiments, for example in embodiments in which the fan diameteris in the range from 330 to 380 cm, the radial bending stiffness of thepin 33 may be in the range from 9.0×10⁸ to 3.0×10⁹ N/m and optionallymay be in the range from 1.0×10⁹ to 2.0×10⁹ N/m (and optionally may beequal to 1.54×10⁹ N/m).

The skilled person would appreciate that the desired pin stiffness maybe obtained in various different ways, for example by adjusting one ormore of pin material(s) (often steels) and gearbox and/or pin geometryas appropriate. For a given material, the stiffness may be a function ofpin diameter and whether the pin is solid, or hollow, for example. Byway of further example, the gear ratio of the gearbox 30 may be adjustedand the planet gear size may change accordingly, for example allowingfor a larger diameter pin 33 for a larger planet gear 32, so achieving ahigher stiffness.

Planet Pin Tilt Stiffness

Each pin 33 (including its soft connections 31 a, 31 b, if present) ismodelled as a free body rigidly mounted on one carrier plate 34 b(illustrated by the diagonal lines on the carrier plate 34 b as shown inFIG. 18) to calculate the tilt stiffness. A moment is then applied atthe axial centre point of the pin 33. The axis of the moment isperpendicular to the engine axis 9. Two points of the pin 33 areselected for measuring tilt stiffness: a central point at the axialposition of the axial centre point of the pin 33, and a rearward pointat the axial position of the rigid connection to the rearward plate 34b. The rearward plate 34 b is held to be rigid and non-rotating, asillustrated by the diagonal lines in FIGS. 17 and 18.

In response to the applied moment, M, which is an anticlockwise momentin the example shown, but could be a clockwise moment in other examples,the pin 33 bends through an angle θ, resulting in an arc displacement 6of the central point (the point of application of the moment).

If the pin 33 is asymmetric, a second measurement may be taken modellingthe pin 33 as a free body rigidly mounted on the other carrier plate 34a. Two points of the pin 33 are selected for this measurement: a centralpoint at the axial position of the axial centre point of the pin 33, anda forward point at the axial position of the rigid connection to theforward plate 34 a. The forward plate 34 a is held to be rigid andnon-rotating. An average of the two tilt stiffness values may then betaken.

An effective linear tilt stiffness can therefore be defined for the pin33 as described above. The radius, r, chosen for the definition of theeffective linear tilt stiffness is that of a circle centred on a pointon the surface of the plate 34 b held to be rigid and passing through anoriginal axial centrepoint of the pin 33 (the point of application ofthe moment) and the same point post-deformation. The arc displacement 6is equal to rθ.

The same approach may be used for any design of pin 33.

In various embodiments, the tilt stiffness of the pin 33 is greater thanor equal to 4.00×10⁶ Nm/rad, and optionally greater than or equal to8.7×10⁶ Nm/rad. In some embodiments, for example in embodiments in whichthe fan diameter is in the range from 240 to 280 cm, the tilt stiffnessof the pin 33 may be greater than or equal to 8.7×10⁶ Nm/rad, andoptionally may be greater than or equal to 9.8×10⁶ Nm/rad (andoptionally may be equal to 1.02×10⁷ Nm/rad). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the tilt stiffness of the pin 33 may be greater than orequal to 1.4×10⁷ Nm/rad and optionally may be greater than or equal to2.5×10⁷ Nm/rad (and optionally may be equal to 3.14×10⁷ Nm/rad).

In various embodiments, the tilt stiffness of the pin 33 is in the rangefrom 4.00×10⁶ to 6.30×10⁷ Nm/rad, and optionally in the range from8.7×10⁶ to 4.5×10⁷ Nm/rad. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the tilt stiffness of the pin 33 may be in the range from 8.7×10⁶ to2.1×10⁷ Nm/rad, and optionally may be in the range from 9.8×10⁶ to1.9×10⁷ Nm/rad (and optionally may be equal to 1.02×10⁷ Nm/rad). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the tilt stiffness of the pin 33 may be inthe range from 1.4×10⁷ to 6.3×10⁷ Nm/rad and optionally may be in therange from 2.5×10⁷ to 4.2×10⁷ Nm/rad (and optionally may be equal to3.14×10⁷ Nm/rad).

In order to obtain a desired pin tilt stiffness, similar options mayapply as for planet pin radial bending stiffness described above.

Stiffnesses of Shafts and Supports

Gearbox Input Shaft Torsional Stiffness

In the arrangement being described, the gearbox input shaft 26 a drivesthe sun gear 28. The gearbox input shaft 26 a may therefore be referredto as a sun input shaft 26 a. The gearbox input shaft 26 a may be a suninput shaft 26 a in star arrangements (as well as planetary). Thegearbox input shaft 26 a may also be referred to as a part of the coreshaft 26—a forward portion 26 a of the core shaft 26 provides the inputto the gearbox 30.

The core shaft 26 therefore comprises a gearbox input shaft 26 a, whichrotates with the rest of the core shaft 26 but may have a differentstiffness from the rest of the core shaft. In the arrangement beingdescribed with respect to FIGS. 1 and 2, the core shaft extends betweenthe turbine 19 and the gearbox 30, connecting the turbine 19 to thecompressor 14, and the turbine and compressor to the gearbox 30. Arearward portion 26 b of the core shaft 26 extends between the turbine19 and the compressor 14, connecting the turbine to the compressor. Aforward portion 26 a extends between the compressor 14 and the gearbox,connecting the turbine and compressor to the gearbox 30. As this forwardportion provides the torque to the gearbox 30, it is referred to as thegearbox input shaft. In the arrangement shown, a bearing 26 c is presenton the core shaft 26 at or near the axial position at which the rearwardportion 26 b meets the gearbox input shaft 26 a.

In some gearboxes 30, the planet carrier 34 may be driven by the coreshaft 26, and more specifically by the gearbox input shaft 26 a, forexample—in such embodiments, the gearbox input shaft 26 a may not be asun input shaft 26. However, this may make mounting of the sun gear 28more difficult.

In the described arrangement, the core shaft 26 is divided into twosections as shown in FIG. 19; a first section 26 a (the gearbox inputshaft) extending from the gearbox 30 and connected to the sun gear 28,and a second section 26 b extending rearwardly from the first sectionand connected to the turbine 19. In the described arrangement, the firstsection 26 a is designed to have a lower stiffness than the secondsection 26 b—the gearbox input shaft 26 a may therefore provide a softmounting for the sun gear 28 whilst maintaining rigidity elsewhere inthe engine 10. In the described arrangement, the second section 26 b isdesigned to be effectively rigid (as compared to the stiffness of thefirst section 26 a)—a torsional stiffness of the core shaft 26 maytherefore be effectively equal to the torsional stiffness of the gearboxinput shaft portion thereof. The second section 26 b connecting theturbine and the compressor may be referred to as the turbine shaft 26 b.The turbine shaft 26 b is arranged to transmit the torsional loads todrive the compressor and the gearbox 30, as well as the compressor andturbine axial loads.

In alternative embodiments, the core shaft 26 may not be divided intosections of different stiffness, and may instead have a constantstiffness. In alternative or additional embodiments, the core shaft 26may be divided into a larger number of sections.

The core shaft 26 is mounted using a bearing 26 c—the bearing 26 c isthe first bearing on the core shaft 26 axially downstream of the gearbox30. In the described arrangement, the bearing 26 c is on the secondsection 26 b of the shaft 26—in other embodiments, it may be on adifferent, or on the only, shaft section. The stiffnesses of the gearboxinput shaft 26 a are measured holding the bearing 26 c rigid, and takingthe connection of the bearing 26 c to the rest of the core shaft 26 b asrigid, such that only the stiffnesses of the first portion 26 a areconsidered. For the purpose of determining torsional stiffness, thegearbox input shaft 26 a is considered to be free at the end to whichthe applied torque τ is applied.

Gearbox input shaft torsional stiffness is a measure of the resistanceof the shaft 26 a to an applied torque, τ, as illustrated in FIG. 20. Itmay be described as resistance to twisting, or winding, of the shaft 26a. The axis of the moment is parallel to the engine axis 9. Thediagonally lined box 402 at the location of the bearing 26 c of theshaft 26 a is shown to indicate the connection to the bearing 26 c/theshaft 26 at the position of the bearing as being treated as rigid andnon-rotating (as for a cantilever beam mounting). The shaft 26 a isotherwise treated as a free body (the sun gear-planet gear meshstiffness is not included).

A torque, τ, is applied to the shaft 26 a (at the forward position—theposition of the axial mid-point of the sun gear 28) and causes arotational deformation, θ (e.g. twist) along the length of the shaft 26a. θ is measured at the position of application of the torque. Asdescribed above, the core shaft 26 is held to be non-rotating at thelocation of the bearing 26 c, such that the value of the twist increasesfrom zero to θ along the length of the first shaft portion 26 a. Theangle through which a point on the shaft circumference at the forwardposition moves is θ, where θ is the angle measured in radians. r is theradius of the shaft 26 a. In embodiments in which the gearbox inputshaft 26 varies in radius, the radius of the shaft 26 a at the interfaceto the sun gear 28 may be used as the radius r to calculate theeffective linear torsional stiffness (i.e. the radius at the forward endof the shaft for the embodiment shown, where the torque is applied). Aneffective linear torsional stiffness can therefore be defined for thegearbox input shaft 26 a as described above.

In the embodiment shown, the position of the axial mid-point of the sungear 28 is also at or near the forward end of the shaft 26. Inalternative embodiments, the shaft 26 may extend further forward of thesun gear 28; the forward position used for the application of thetorque, force or moment is still taken to be the position of the axialmid-point of the sun gear 28 in such embodiments.

In various embodiments, the torsional stiffness of the gearbox inputshaft 26 a is greater than or equal to 1.4×10⁶ Nm/radian, and optionallygreater than or equal to 1.6×10⁶ Nm/radian. In some embodiments, forexample in embodiments in which the fan diameter is in the range from240 to 280 cm, the torsional stiffness of the gearbox input shaft may begreater than or equal to 1.4×10⁶ Nm/radian or 1.8×10⁶ Nm/radian. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the torsional stiffness of the gearboxinput shaft may be greater than or equal to 3×10⁶ Nm/radian or 5×10⁶Nm/radian.

In various embodiments, the torsional stiffness of the gearbox inputshaft 26 a is in the range from 1.4×10⁶ to 2.5×10⁸ Nm/radian, andoptionally in the range from 1.6×10⁶ to 2.5×10⁷ Nm/radian. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the torsional stiffness of the gearboxinput shaft may be in the range from 1.4×10⁶ to 2.0×10⁷ Nm/radian, andoptionally may be in the range from 1.8×10⁶ to 3×10⁶ Nm/radian (andoptionally may be equal to 2.0×10⁶ Nm/radian). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the torsional stiffness of the gearbox input shaft may bein the range from 3×10⁶ to 1×10⁸ Nm/radian and optionally may be in therange from 5×10⁶ to 6×10⁶ Nm/radian (and optionally may be equal to5.7×10⁶ Nm/radian).

In various embodiments, the effective linear torsional stiffness of thegearbox input shaft 26 a is greater than or equal to 4.0×10⁸ N/m, andoptionally greater than or equal to 4.3×10⁸ N/m. In some embodiments,for example in embodiments in which the fan diameter is in the rangefrom 240 to 280 cm, the effective linear torsional stiffness of thegearbox input shaft may be greater than or equal to 4.0×10⁸ N/m or4.4×10⁸ N/m. In some embodiments, for example in embodiments in whichthe fan diameter is in the range from 330 to 380 cm, effective lineartorsional stiffness of the gearbox input shaft may be greater than orequal to 4.3×10⁸ N/m or 6.8×10⁸ N/m.

In various embodiments, the effective linear torsional stiffness of thegearbox input shaft is in the range 4.0×10⁸ to 3.0×10¹⁰ N/m, andoptionally in the range from 4.3×10⁸ to 9.0×10⁹ N/m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the effective linear torsional stiffnessof the gearbox input shaft may be in the range from 4.0×10⁸ to 1.5×10¹⁰N/m, and optionally may be in the range from 4.4×10⁸ to 5.4×10⁹ N/m (andoptionally may be equal to 4.9×10⁸ N/m, and optionally 4.92×10⁸ N/m). Insome embodiments, for example in embodiments in which the fan diameteris in the range from 330 to 380 cm, the effective linear torsionalstiffness of the gearbox input shaft may be in the range from 4.3×10⁸ to3.0×10¹⁰ N/m and optionally may be in the range from 5.0×10⁸ to 8.0×10¹⁰N/m (and optionally may be equal to 6.8×10⁸ N/m, and optionally 6.84×10⁸N/m).

One or more of gearbox input shaft 26 a material(s), diameter andstructure (e.g. hollow or solid, wall thickness) may be adjusted toachieve a stiffness within the desired range.

Fan Shaft Torsional Stiffness

The fan shaft 36 is defined as the torque transfer component thatextends from the output of the gearbox 30 to the fan input. It thereforeincludes part or all of any gearbox output shaft and fan input shaftthat may be provided between those points. For the purposes of definingthe stiffness of the fan shaft 36, it is considered to extend between afan input position and a gearbox output position, and to include all ofthe torque transfer components between those points. It does nottherefore include any components of the gearbox 30 (e.g. the planetcarrier 34 or any connecting plate coupled to it) which transmitdiscrete forces, rather than the fan shaft torque. The gearbox outputposition therefore may be defined as the point of connection between thefan shaft 36 and the gearbox 30. The fan input position, Y, may bedefined as the point of connection between the fan shaft 36 and the fan.The torsional stiffness of the fan shaft 36 is measured between theforward and rearward ends of the fan shaft; the forward end being theinterface (Y) with the fan 23 and the rearward end the interface (X)with the gearbox 30.

Fan shaft torsional stiffness is a measure of the resistance of theshaft 36 to an applied torque, τ, as illustrated in FIG. 25. It may bedescribed as resistance to twisting, or winding, of the shaft 36. Theaxis of the moment is parallel to the engine axis 9.

Referring to FIGS. 23 and 25, where the gearbox 30 is a star gearbox,the gearbox output position (X) is defined as the point of connection702 between the ring gear 38 and the fan shaft 36. More specifically, itis the point of connection to the annulus of the ring gear 38 (with anyconnection component extending from the outer surface of the annulusbeing considered to be part of the ring gear). Where the point ofconnection is formed by an interface extending in a direction having anaxial component, the point of connection is considered to be the axialcentreline (X) of that interface as illustrated in FIG. 25. The fanshaft 36 includes all torque transmitting components up to the point ofconnection 702 with the ring gear 38. It therefore includes any flexibleportions or linkages 704 of the fan shaft 36 that may be provided, andany connection(s) 706 (e.g. spline connections) between them.

Where the gearbox 30 is in a planetary configuration, the gearbox outputposition is again defined as the point of connection between the fanshaft 36 and the gearbox 30. An example of this is illustrated in FIG.24, which shows a carrier comprising a forward plate 34 a and rearwardplate 34 b, with a plurality of pins 33 extending between them and onwhich the planet gears are mounted. The fan shaft 36 is connected to theforward plate 34 a via a spline connection 708. In an embodiment such asthis, the gearbox output position, X, is taken as any point on theinterface between the fan shaft 36 and the forward plate 34 a. Theforward plate 34 a is considered to transmit discrete forces, ratherthan a single torque, and so is taken to be part of the gearbox 30rather than the fan shaft. FIG. 24 shows only one example of a type ofconnection between the fan shaft and planet carrier 34. In embodimentshaving different connection arrangements, the gearbox output position isstill taken to be at the interface between components transmitting atorque (i.e. that are part of the fan shaft) and those transmittingdiscrete forces (e.g. that are part of the gearbox). The splineconnection 708 is only one example of a connection that may be formedbetween the fan shaft 36 and gearbox 30 (i.e. between the fan shaft andthe forward plate 34 b in the presently described embodiment). In otherembodiments, the interface which forms the gearbox output position maybe formed by, for example, a curvic connection, a bolted joint or othertoothed or mechanically dogged arrangement.

The fan input position, Y, is defined as a point on the fan shaft 36 atthe axial midpoint of the interface between the fan 23 and the fan shaft36. In the presently described embodiment, the fan 23 comprises asupport arm 23 a arranged to connect the fan 23 to the fan shaft 36. Thesupport arm 23 a is connected to the fan shaft by a spline coupling 36 athat extends along the length of a portion of the fan shaft 36. The faninput position is defined as the axial midpoint of the spline couplingas indicated by axis Y in FIG. 25. The spline coupling shown in FIG. 25is only one example of a coupling that may form the interface betweenthe fan and fan shaft. In other embodiments, for example, a curvicconnection, a bolted joint or other toothed or mechanically doggedarrangement may be used. The fan input position, Y, may be unaffected bygearbox type.

The fan shaft 36 has a degree of flexibility characterised in part byits torsional stiffness. For the purpose of determining torsionalstiffness, the fan shaft 36 is considered to be free at the end to whichthe applied torque, τ, is applied.

The diagonally lined ring gear 38 in FIG. 25 indicates the ring gear 38being treated as rigid and non-rotating for the purpose of assessingtorsional stiffness of the fan shaft 36. A torque, τ, is applied to theshaft 36 at the fan input position, Y, and causes a rotationaldeformation, θ (e.g. twist) along the length of the shaft 36. The anglethrough which a point on the shaft circumference at the fan inputposition moves is θ, where θ is the angle measured in radians. Theradius, r, of the fan shaft 36 may be used to determine an effectivelinear torsional stiffness. In embodiments in which the fan shaft 36varies in radius, such as the embodiment being described, the radius ofthe shaft 36 at the fan input position may be used as the radius r (i.e.the radius at the forward end of the shaft for the embodiment shown). Aneffective linear torsional stiffness can therefore be defined for thefan shaft 36 as described above.

In various embodiments, the torsional stiffness of the fan shaft 36 isequal to or greater than 1.3×10⁷ Nm/rad, and optionally equal to orgreater than 1.4×10⁷ Nm/rad. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280 cmthe torsional stiffness of the fan shaft 36 may equal to or greater than1.3×10⁷ Nm/radian, and optionally may be equal to or greater than1.4×10⁷ Nm/radian (and optionally may be equal to 1.8×10⁷ Nm/radian). Insome embodiments, for example in embodiments in which the fan diameteris in the range from 330 to 380 cm, the torsional stiffness of the fanshaft 36 may be equal to or greater than 2.5×10⁷ Nm/radian andoptionally may be equal to or greater than 3.5×10⁷ Nm/radian (andoptionally may be equal to 5.2×10⁷ Nm/radian).

In various embodiments, the torsional stiffness of the fan shaft 36 isin the range from 1.3×10⁷ to 2.5×10⁹ Nm/rad, and optionally in the rangefrom 1.4×10⁷ to 3.0×10⁸ Nm/rad. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280 cmthe torsional stiffness of the fan shaft 36 may be in the range from1.3×10⁷ to 2.0×10⁸ Nm/radian, and optionally may be in the range from1.3×10⁷ or 1.4×10⁷ to 2.3×10⁷ Nm/radian (and optionally may be equal to1.8×10⁷ Nm/radian). In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 330 to 380 cm, the torsionalstiffness of the fan shaft 36 may be in the range from 2.5×10⁷ to2.5×10⁹ Nm/radian and optionally may be in the range from 3.5×10⁷ to7.5×10⁷ Nm/radian (and optionally may be equal to 5.2×10⁷ Nm/radian).

In various embodiments, the effective linear torsional stiffness of thefan shaft 36 may be greater than 1.2×10⁹ N/m, and optionally greaterthan 1.35×10⁹ N/m. In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 240 to 280 cm, the effectivelinear torsional stiffness of the fan shaft 36 may be greater than1.2×10⁹ N/m, and optionally may be greater than 1.3×10⁹ Nm/radian (andoptionally may be equal to 1.5×10⁹ N/m). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the effective linear torsional stiffness of the fan shaft36 may be greater than 1.5×10⁹ N/m and optionally may be greater than1.8×10⁹ Nm/radian (and optionally may be equal to 2.1×10⁹ N/m).

In various embodiments, the effective linear torsional stiffness of thefan shaft 36 is in the range from 1.2×10⁹ to 2.0×10¹⁰ N/m, andoptionally in the range from 1.35×10⁹ to 1.0×10¹⁰ N/m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the effective linear torsional stiffnessof the fan shaft 36 may be in the range from 1.2×10⁹ to 1.5×10¹⁰ N/m,and optionally may be in the range from 1.3×10⁹ to 2.3×10⁹ Nm/radian(and optionally may be equal to 1.5×10⁹ N/m). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the effective linear torsional stiffness of the fan shaft36 may be in the range from 1.5×10⁹ to 2.0×10¹⁰ N/m and optionally maybe in the range from 1.8×10⁹ to 3.5×10⁹ Nm/radian (and optionally may beequal to 2.1×10⁹ N/m).

One or more of fan shaft 36 material(s), diameter and structure (e.g.hollow or solid, wall thickness) may be adjusted to achieve a stiffnesswithin the desired range.

In various embodiments, including the arrangements being described, e.g.with respect to FIGS. 21 and 22, a mounting structure for the fan shaft36 comprises a fan shaft support structure 504. The fan shaft supportstructure comprises two bearings—a first bearing 506 a and a secondbearing 506 b—via which it is coupled to the fan shaft 36. The bearings506 a, 506 a are spaced apart along the axial length of the fan shaft36. In the described embodiment, both bearings 506 a, 506 b are providedat positions that are forward of the gearbox 30. In other embodiments,one of the two bearing 506 a, 506 b used to support the fan shaft 36 maybe located at a position rearward of the gearbox 30. In yet otherembodiments, more than two bearings may be provided as part of the fanshaft support structure.

Gearbox Support Torsional Stiffness

An exemplary embodiment of the gas turbine engine is shown in FIG. 22,which provides an enlarged view of a region of the engine core 11 aroundthe gearbox 30. The same reference numbers have been used for componentscorresponding to those shown in FIGS. 1 to 3. In the arrangement shownin FIG. 22 the gearbox 30 has a star arrangement, in which the ring gear38 is coupled to the fan shaft 36 and the carrier 34 is held in a fixedposition relative to the static structure 24 (also referred to as thestationary supporting structure) of the engine core. As noted elsewhereherein, all features and characteristics described herein may apply to astar gearbox and a planetary gearbox, unless explicitly statedotherwise.

The engine core 11 comprises a gearbox support 40 (corresponding to thelinkage described with reference to FIG. 2) arranged to at leastpartially support the gearbox 30 in a fixed position within the engine10. The gearbox support is coupled at a first end to the stationarysupporting structure 24 which extends across the core duct 20 carryingthe core airflow A as illustrated in FIG. 22. In the presently describedarrangement, the stationary support structure 24 is or comprises anengine section stator (ESS) that acts as both a structural component toprovide a stationary mounting for core components such as the gearboxsupport 40, and as a guide vane provided to direct airflow from the fan23. In other embodiments, the stationary supporting structure 24 maycomprise a strut extending across the core gas flow path and a separatestator vane provided to direct airflow. In the present embodiment, thegearbox support 40 is coupled at a second end to the planet carrier 34.The gearbox support 40 therefore acts against rotation of the planetcarrier 34 relative to the static structure 24 of the engine core. Inembodiments where the gearbox 30 is in a planetary arrangement, thegearbox support 40 is coupled to the ring gear 38 so as to resist itsrotation relative to the static structure 24 of the engine core.

The gearbox support 40 is defined between the point at which it connectsto the gearbox (e.g. to the planet carrier in the present embodimentwith a star gearbox 30, or to the ring gear 38 in planetary embodiments)and a point at which it connects to the stationary supporting structure24. The gearbox support may be formed by any number of separatecomponents providing a coupling between those two points. The gearboxsupport 40 connects to the gearbox 30, and more specifically to thestatic gear or gearset—i.e. to the ring gear 38 of a planetary gearboxor the planet carrier 34/planet gear set 32 of a star gearbox. Thegearbox support 40 has a degree of flexibility. Gearbox supporttorsional stiffness is a measure of the resistance of the support 40 toan applied torque, τ, as illustrated in FIG. 21. It may be described asresistance to twisting, or winding, of the support 40. The axis of themoment is parallel to the engine axis 9. The cross-hatching of thestationary support structure 24 is provided to indicate the connectionto the support 40 being treated as rigid and non-rotating.

For a star gearbox 30, the torsional stiffness of the gearbox support 40is defined between a circle 902 passing through the centre of eachplanet gear 32 of the planetary gear set (i.e. through the longitudinalaxis of each pin 33) and the interface to the stationary supportstructure 24, which is treated as fixed. The torsional load is appliedat the planet carrier 34, and reacted at the stationary supportstructure 24.

For a planetary gearbox 30, the torsional stiffness of the gearboxsupport 40 is defined between the pitch circle diameter (PCD) of thering gear 38, and the interface to the stationary support structure 24,which is treated as fixed. The torsional load is applied at the ringgear 38, and reacted at the stationary support structure 24. A torque,τ, is applied to the teeth of the ring gear 38 and causes a rotationaldeformation, θ (e.g. twist) of the support 40. The angle through which apoint on the PCD moves is θ, where θ is the angle measured in radians. Aradius, r, may be defined as the radius of the ring gear 38 (i.e. halfof the PCD of the ring gear). An effective linear torsional stiffnesscan therefore be defined for the gearbox support 40 for a planetarygearbox 30 as described above using the radius r=PCD/2.

The pitch circle of a gear is an imaginary circle that rolls withoutslipping with the pitch circle of any other gear with which the firstgear is meshed. The pitch circle passes through the points where theteeth of two gears meet as the meshed gears rotate—the pitch circle of agear generally passes through a mid-point of the length of the teeth ofthe gear. The PCD can be roughly estimated by taking the average of thediameter between tips of the gear teeth and the diameter between basesof the gear teeth. In various embodiments the PCD of the ring gear 38,which may also be thought of as a diameter of the gearbox 30, may begreater than or equal to 0.55 m, and optionally greater than or equal to0.57 m. In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the gearbox diameter may begreater than or equal to 0.75 m.

In various embodiments the diameter of the gearbox 30 may be in therange from 0.55 m to 1.2 m, and optionally in the range from 0.57 to 1.0m. In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm the gearbox diameter may bein the range from 0.55 to 0.70 m, and optionally may be in the rangefrom 0.58 to 0.65 m (and optionally may be equal to 0.61 m). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the gearbox diameter may be in the rangefrom 0.75 to 1.0 m, and optionally from 0.8 to 0.9 m (and optionally maybe equal to 0.87 m).

Correspondingly, an effective linear torsional stiffness can thereforebe defined for the gearbox support 40 for a star gearbox 30 as describedabove using the radius r of the circle 902 passing through thelongitudinal axis of each pin 33 on the carrier 34. The diameter of thiscircle 902 may be described as a PCD of the planetary gear set, or a pinPCD, so providing r=PCD/2 as for the planetary gearbox example. The pinPCD may be in the range from 0.38 to 0.65 m, for example being equal to0.4 m or 0.55 m.

In various embodiments, the torsional stiffness of the gearbox support40 is greater than or equal to 4.2×10⁷ Nm/rad, and optionally greaterthan or equal to 4.8×10⁷ Nm/rad. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280 cmthe torsional stiffness of the gearbox support 40 may be greater than orequal to 4.2×10⁷ Nm/rad, and optionally may be greater than or equal to5×10⁷ Nm/rad (and optionally may be equal to 6.1×10⁷ Nm/rad). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the torsional stiffness of the gearboxsupport 40 may be greater than or equal to 7.0×10⁷ Nm/rad, andoptionally may be greater than or equal to 9×10⁷ Nm/rad (and optionallymay be equal to 1.8×10⁸ Nm/rad).

In various embodiments, the torsional stiffness of the gearbox support40 is in the range from 4.2×10⁷ to 1.0×10¹⁰ Nm/rad, and optionally inthe range from 4.8×10⁷ to 1.0×10⁹ Nm/rad. In some embodiments, forexample in embodiments in which the fan diameter is in the range from240 to 280 cm the torsional stiffness of the gearbox support 40 may bein the range from 4.2×10⁷ to 6.0×10⁸ Nm/rad, and optionally may be inthe range from 5×10⁷ to 7×10⁷ Nm/rad (and optionally may be equal to6.1×10⁷ Nm/rad). In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 330 to 380 cm, the torsionalstiffness of the gearbox support 40 may be in the range from 7.0×10⁷ to1.0×10¹⁰ Nm/rad, and optionally may be in the range from 9×10⁷ to 4×10⁸Nm/rad (and optionally may be equal to 1.8×10⁸ Nm/rad).

In various embodiments, the effective linear torsional stiffness of thegearbox support 40 is greater than 7.1×10⁸ N/m, and optionally greaterthan 8.4×10⁸ N/m. In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 240 to 280 cm the effectivelinear torsional stiffness of the gearbox support 40 may be greater than7.1×10⁸ N/m, and optionally may be greater than 8×10⁸ N/m (andoptionally may be equal to 9.2×10⁸ N/m). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the effective linear torsional stiffness of the gearboxsupport 40 may be greater than 9.0×10⁸ N/m, and optionally may begreater than 9.6×10⁸ N/m (and optionally may be equal to 1.2×10⁹ N/m).

In various embodiments, the effective linear torsional stiffness of thegearbox support 40 is in the range from 7.1×10⁸ to 6.0×10¹⁰ N/m, andoptionally in the range from 8.4×10⁸ to 3.0×10¹⁰ N/m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm the effective linear torsional stiffness ofthe gearbox support 40 may be in the range from 7.1×10⁸ to 5.0×10¹⁰ N/m,and optionally may be in the range from 8×10⁸ to 1×10⁹ N/m (andoptionally may be equal to 9.2×10⁸ N/m). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the effective linear torsional stiffness of the gearboxsupport 40 may be in the range from 9.0×10⁸ to 6.0×10¹⁰ N/m, andoptionally may be in the range from 9.0×10⁸ to 2.0×10⁹ N/m (andoptionally may be equal to 1.2×10⁹ N/m).

The skilled person would appreciate that the stiffness of the gearboxsupport 40 may be defined in the same way for embodiments with differentepicyclic gearboxes, e.g. planetary gearboxes.

One or more of gearbox support 40 geometry, materials, and connectiontype for the connection to the stationary support structure 24 may beselected or adjusted as appropriate to obtain the desired stiffness.

The inventor has discovered that particular ratios of the parametersdefined above have significant impact on gearbox performance. Inparticular, one, some or all of the below may apply to any embodiment:

In various embodiments, a radial to torsional carrier stiffness ratioof:

$\frac{{the}\mspace{14mu}{radial}\mspace{14mu}{bending}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}{\begin{matrix}{{{the}\mspace{14mu}{effective}\mspace{14mu}{linear}\mspace{14mu}{torsional}}\mspace{14mu}} \\{{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}\end{matrix}}$

-   -   is greater than or equal to 0.030, and optionally in the range        from 0.030 to 2.0.

In various embodiments, the radial to torsional carrier stiffness ratiois in the range from 3.0×10⁻² to 2.0×10⁰ (i.e. from 0.030 to 2.0), andoptionally in the range from 6.0×10⁻² to 1.0. In some embodiments, theradial to torsional carrier stiffness ratio may be in the range from6.0×10⁻² to 3.0×10⁻¹, and optionally may be in the range from 0.18 to0.19 (and optionally may be equal to 0.18). In some embodiments, theradial to torsional carrier stiffness ratio may be in the range from0.30 to 2.0. In alternative such embodiments, the radial to torsionalcarrier stiffness ratio may be in the range from 0.14 to 0.8, andoptionally may be in the range from 0.14 to 0.19 (and optionally may beequal to 0.14).

In various embodiments, a product of the components of the radial totorsional carrier stiffness ratio, i.e. the radial bending stiffness ofthe planet carrier 34 multiplied by the effective linear torsionalstiffness of the planet carrier 34, may be calculated. The value of thisproduct, in various embodiments, may be greater than or equal to5.0×10¹⁸ N²m⁻², and optionally less than 1.3×10²⁴ N²m⁻², and optionallymay be greater than or equal to 1.6×10¹⁹ N²m⁻², and optionally less than1.3×10²² N² m⁻². In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 240 to 280 cm, the productvalue may be greater than or equal to 1.6×10¹⁹ N² m⁻², and optionallyless than 1.3×10²² N² m⁻². In some embodiments, for example inembodiments in which the fan diameter is in the range from 330 to 380cm, the product value may be greater than or equal to 3.0×10¹⁹ N² m⁻²,and optionally less than 1.3×10²³ N² m⁻².

In various embodiments, a tilt to torsional carrier stiffness ratio of:

$\frac{{the}\mspace{14mu}{tilt}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}{{the}\mspace{20mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}$

is less than or equal to 20, and optionally less than or equal to 7.3.

In various embodiments, the tilt to torsional carrier stiffness ratio isin the range from 7.00×10⁻¹ to 2.0×10¹ (i.e. from 0.7 to 20), andoptionally in the range from 0.7 to 7.3. In some embodiments, forexample in embodiments in which the fan diameter is in the range from240 to 280 cm the tilt to torsional carrier stiffness ratio may be lessthan or equal to 8.0, optionally in the range from 2.5 to 8.0, andfurther optionally may be in the range from 4 to 7 (and optionally maybe equal to 5.60). In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 330 to 380 cm, the tilt totorsional carrier stiffness ratio may be less than or equal to 7.9,optionally in the range from 1.5 to 7.9, and further optionally may bein the range from 1.8 to 5.2 (and optionally may be equal to 2.63).

In various embodiments, a product of the components of the tilt totorsional carrier stiffness ratio, i.e. the tilt stiffness of the planetcarrier (34) multiplied by the torsional stiffness of the planet carrier(34), may be calculated. The value of this product, in variousembodiments, may be greater than or equal to 1.0×10¹ N²m²rad⁻², andoptionally less than 2.8×10²² N²m²rad⁻², and optionally may be greaterthan or equal to 5.1×10¹ N²m²rad⁻², and optionally less than 3.0×10²¹N²m²rad⁻². In some embodiments, for example in embodiments in which thefan diameter is in the range from 240 to 280 cm, the product value maybe greater than or equal to 5.1×10¹⁷ N²m²rad⁻², and optionally less than3.0×10²⁰ N²m²rad⁻². In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 330 to 380 cm, the productvalue may be greater than or equal to 1.0×10¹⁸ N²m²rad⁻², and optionallyless than 3.1×10²¹ N²m²rad⁻².

In various embodiments, a carrier to gearbox input shaft torsionalstiffness ratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{gearbox}\mspace{14mu}{input}\mspace{14mu}{shaft}\mspace{11mu}\left( {26a} \right)}$

-   -   is greater than or equal to 7.0×10¹.

In various embodiments, the carrier to gearbox input shaft torsionalstiffness ratio may be greater than or equal to 7.5×10¹. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm the carrier to gearbox input shafttorsional stiffness ratio may greater than or equal to 7.3×10¹, andoptionally may be greater than or equal to 9.5×10¹ or 14.0×10¹ (andoptionally may be equal to 152). In some embodiments, for example inembodiments in which the fan diameter is in the range from 330 to 380cm, the carrier to gearbox input shaft torsional stiffness ratio may begreater than or equal to 1.0×10², and optionally may be greater than orequal to 1.5×10² (and optionally may be equal to 2.0×10²). In variousembodiments, the carrier to gearbox input shaft torsional stiffnessratio may be greater than or equal to 1.4×10², and optionally in therange from 1.4×10² and to 5.4×10².

In various embodiments, the carrier to gearbox input shaft torsionalstiffness ratio may be in the range from 7×10¹ to 5×10³ (i.e. 70 to5000), and optionally from 7.5×10¹ to 3.0×10³. In some embodiments, forexample in embodiments in which the fan diameter is in the range from240 to 280 cm the carrier to gearbox input shaft torsional stiffnessratio may be in the range from 7.3×10¹ to 1.0×10³, and optionally may bein the range from 9.5×10¹ or 14.0×10¹ to 3.0×10² (and optionally may beequal to 152). In some embodiments, for example in embodiments in whichthe fan diameter is in the range from 330 to 380 cm, the carrier togearbox input shaft torsional stiffness ratio may be in the range from1.0×10² to 5.0×10³, and optionally may be in the range from 1.5×10² to2.7×10² (and optionally may be equal to 2.0×10²).

In various embodiments, a product of the components of the carrier togearbox input shaft torsional stiffness ratio, i.e. the torsionalstiffness of the planet carrier (34) multiplied by the torsionalstiffness of the gearbox input shaft (26 a), may be calculated. Thevalue of this product, in various embodiments, may be greater than orequal to 1.5×10¹⁴ N²m²rad⁻², and optionally less than 1.0×10¹ N²m²rad⁻²,and optionally may be greater than or equal to 2.2×10¹⁴ N²m²rad⁻², andoptionally less than 5.0×10¹⁶ N²m²rad⁻². In some embodiments, forexample in embodiments in which the fan diameter is in the range from240 to 280 cm, the product value may be greater than or equal to1.5×10¹⁴ N²m²rad⁻², and optionally less than 1.0×10¹⁶ N²m²rad⁻². In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the product value may be greater than orequal to 3.0×10¹ N²m²rad⁻², and optionally less than 1.0×10¹ N²m²rad⁻².

In various embodiments, a carrier to gearbox support torsional stiffnessratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{gearbox}\mspace{14mu}{support}\mspace{11mu}(40)}$

-   -   is greater than or equal to 2.3.

In various embodiments, the carrier to gearbox support torsionalstiffness ratio may be greater than or equal to 2.3, and optionallygreater than or equal to 2.6. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280 cmcarrier to gearbox support torsional stiffness ratio may be greater thanor equal to 2.3, and optionally may be greater than or equal to 2.5 (andoptionally may be equal to 4.8). In some embodiments, for example inembodiments in which the fan diameter is in the range from 330 to 380cm, the carrier to gearbox support torsional stiffness ratio may begreater than or equal to 3.5, and optionally may be greater than orequal to 4 (and optionally may be equal to 6.5). In various embodiments,the carrier to gearbox support torsional stiffness ratio may be greaterthan or equal to 4.4, and optionally in the range from 4.4 or 4.5 to15.5.

In various embodiments, the carrier to gearbox support torsionalstiffness ratio may be in the range from 2.3×10⁰ to 3.0×10² (i.e. 2.3 to300), and optionally from 2.6 to 50. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the carrier to gearbox support torsional stiffness ratio may be inthe range from 2.3 to 30, and optionally may be in the range from 2.5 to5.5 or from 4.3 to 5.5 (and optionally may be equal to 4.8). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the carrier to gearbox support torsionalstiffness ratio may be in the range from 3.5 to 300, and optionally maybe in the range from 4 to 15 (and optionally may be equal to 6.5).

In various embodiments, a product of the components of the carrier togearbox support torsional stiffness ratio i.e. the torsional stiffnessof the planet carrier 34 multiplied by the torsional stiffness of thegearbox support 40, may be calculated. The value of this product, invarious embodiments, may be greater than or equal to 5.0×10¹⁵ N²m²rad⁻², and optionally less than 1.0×10¹⁹ N² m²rad⁻², and optionallymay be greater than or equal to 8.0×10¹⁵ N² m²rad⁻², and optionally lessthan 2.0×10¹⁸ N² m²rad⁻². In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the product value may be greater than or equal to 5.0×10¹⁵ N²m²rad⁻², and optionally less than to 1.2×10¹⁷ N² m²rad⁻². In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the product value may be greater than orequal to 1.0×10¹⁷ N² m²rad⁻², and optionally less than 1.0×10¹⁹ N²m²rad⁻².

In various embodiments, a carrier to fan shaft stiffness ratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{20mu}{fan}\mspace{20mu}{shaft}\mspace{11mu}(36)}$

-   -   is greater than or equal to 8.

In various embodiments, the carrier to fan shaft stiffness ratio may begreater than or equal to 8.0×10⁰ (i.e. 8.0), and optionally greater thanor equal to 9. In some embodiments, for example in embodiments in whichthe fan diameter is in the range from 240 to 280 cm, the carrier to fanshaft stiffness ratio may be greater than or equal to 8, and optionallymay be greater than or equal to 9 or 15.1 (and optionally may be equalto 16.6). In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the carrier to fanshaft stiffness ratio may be greater than or equal to 12, and optionallymay be greater than or equal to 15 or 18 (and optionally may be equal to22.2). In various embodiments, the carrier to fan shaft stiffness ratiomay be greater than or equal to 1.50×10¹, and optionally greater than orequal to 1.6×10¹; the carrier to fan shaft stiffness ratio may besmaller than 8.4×10¹ in such embodiments.

In various embodiments, the carrier to fan shaft stiffness ratio may bein the range from 8.0×10⁰ to 1.1×10³ (i.e. 8.0 to 1100), and optionallyfrom 9 to 1.9×10². In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 240 to 280 cm, the carrierto fan shaft stiffness ratio may be in the range from 8 to 5.0×10², andoptionally may be in the range from 9 to 40 or from 15 or 16 to 40 (andoptionally may be equal to 17). In some embodiments, for example inembodiments in which the fan diameter is in the range from 330 to 380cm, the carrier to fan shaft stiffness ratio may be in the range from 12to 1.1×10³, and optionally may be in the range from 15 or 18 to 55 (andoptionally may be equal to 22).

In various embodiments, a product of the components of the carrier tofan shaft stiffness ratio i.e. the torsional stiffness of the planetcarrier 34 multiplied by the torsional stiffness of the fan shaft 36,may be calculated. The value of this product, in various embodiments,may be greater than or equal to 1.5×10¹⁵ N² m²rad⁻², and optionally lessthan 3.0×10¹⁸ N² m²rad⁻², and optionally may be greater than or equal to2.0×10¹⁵ N²m²rad⁻², and optionally less than 7.0×10¹⁷ N²m²rad⁻². In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the product value may be greater than orequal to 1.5×10¹⁵ N² m²rad⁻², and optionally less than to 1.5×10¹⁷ N²m²rad⁻². In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the product value maybe greater than or equal to 9.0×10¹⁵ N²m²rad⁻², and optionally less than3.0×10¹⁸ N²m²rad⁻².

In various embodiments, a first carrier to pin stiffness ratio of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{effective}\mspace{14mu}{linear}\mspace{20mu}{torsional}}\mspace{14mu}} \\{{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}\end{matrix}}{{the}\mspace{14mu}{radial}\mspace{14mu}{bending}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}\mspace{11mu}(33)}$

-   -   is greater than or equal to 10, and optionally greater than or        equal to 15. In some embodiments, for example in embodiments in        which the fan diameter is in the range from 240 to 280 cm, the        first carrier to pin stiffness ratio shaft may be greater than        or equal to 1.5×10¹, and optionally may be equal to 16.3. In        some embodiments, for example in embodiments in which the fan        diameter is in the range from 330 to 380 cm, the first carrier        to pin stiffness ratio may be greater than or equal to 1.6×10¹        and optionally may be greater than or equal to 16.5 (and        optionally may be equal to 18.7).

In various embodiments, the first carrier to pin stiffness ratio is inthe range from 1.0×10¹ to 4.0×10¹ (i.e. from 10 to 40), and optionallyin the range from 1.5×10¹ to 3.0×10¹. In some embodiments, for examplein embodiments in which the fan diameter is in the range from 240 to 280cm, the first carrier to pin stiffness ratio shaft may be in the rangefrom 1.5×10¹ to 2.5×10¹, and optionally may be in the range from 15 to19 (and optionally may be equal to 16.3). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the first carrier to pin stiffness ratio may be in therange from 1.6×10¹ to 3.5×10¹ and optionally may be in the range from 16or 16.5 to 20 (and optionally may be equal to 18.7).

The product of the effective linear torsional stiffness of the planetcarrier 34 and the radial bending stiffness of each pin 33 may begreater than or equal to 2.1×10¹⁸ N² m⁻², and optionally greater than orequal to 5.8×10¹⁸ N² m⁻². In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the product of the effective linear torsional stiffness of theplanet carrier 34 and the radial bending stiffness of each pin 33 may begreater than or equal to 5.3×10¹⁸ N² m⁻². In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the product of the effective linear torsional stiffnessof the planet carrier 34 and the radial bending stiffness of each pin 33may be greater than or equal to 1.2×10¹⁹ N²m⁻².

The product of the effective linear torsional stiffness of the planetcarrier 34 and the radial bending stiffness of each pin 33 may be in therange from 2.1×10¹⁸ to 3.6×10²⁰ N² m⁻², and optionally in the range from5.8×10¹⁸ to 1.7×10²⁰ N² m⁻². In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the product of the effective linear torsional stiffness of theplanet carrier 34 and the radial bending stiffness of each pin 33 may bein the range from 5.3×10¹⁸ to 4.0×10¹⁹ N²m⁻². In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the product of the effective linear torsional stiffnessof the planet carrier 34 and the radial bending stiffness of each pin 33may be in the range from 1.2×10¹⁹ to 1.7×10²⁰ N² m⁻².

In various embodiments, a second carrier to pin stiffness ratio of:

$\frac{{the}\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}\mspace{11mu}(34)}{{the}\mspace{14mu}{tilt}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{pin}\mspace{11mu}(33)}$

-   -   is greater than or equal to 2.4×10¹, and optionally greater than        or equal to 3.4×10¹. In some embodiments, for example in        embodiments in which the fan diameter is in the range from 240        to 280 cm, the second carrier to pin stiffness ratio may be        greater than or equal to 3.4×10¹, optionally greater than or        equal to 36, and optionally may be equal to 47.5. In some        embodiments, for example in embodiments in which the fan        diameter is in the range from 330 to 380 cm, the second carrier        to pin stiffness ratio may be greater than or equal to 4.0×10¹        and optionally may be greater than or equal to 45 (and        optionally may be equal to 69.1).

In various embodiments, the second carrier to pin stiffness ratio is inthe range from 2.4×10¹ to 1.8×10¹ (i.e. from 24 to 180), and optionallyin the range from 3.4×10¹ to 1.4×10². In some embodiments, for examplein embodiments in which the fan diameter is in the range from 240 to 280cm, the second carrier to pin stiffness ratio shaft may be in the rangefrom 3.4×10¹ to 1.2×10², and optionally may be in the range from 36 to58 (and optionally may be equal to 47.5).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the second carrier to pinstiffness ratio may be in the range from 4.0×10¹ to 1.8×10² andoptionally may be in the range from 45 to 95 (and optionally may beequal to 69.1).

The product of the torsional stiffness of the planet carrier 34 and thetilt stiffness of each pin 33 may be greater than or equal to 1.0×10¹⁵N² m²rad⁻², and optionally greater than or equal to 2.5×10¹⁵ N² m²rad⁻².In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the product of thetorsional stiffness of the planet carrier 34 and the tilt stiffness ofeach pin 33 may be greater than or equal to 2.5×10¹⁵ N² m²rad⁻². In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the product of the torsional stiffness ofthe planet carrier 34 and the tilt stiffness of each pin 33 may begreater than or equal to 1.4×10¹⁶ N²m²rad⁻².

The product of the torsional stiffness of the planet carrier 34 and thetilt stiffness of each pin 33 may be in the range from 1.0×10¹⁵ to4.7×10¹⁷ N² m²rad⁻², and optionally in the range from 2.5×10¹⁵ to2.0×10¹⁷ N²m²rad⁻². In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 240 to 280 cm, the productof the torsional stiffness of the planet carrier 34 and the tiltstiffness of each pin 33 may be in the range from 2.5×10¹⁵ to 3.0×10¹⁶N² m²rad⁻². In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the torsional stiffnessof the planet carrier 34 and the tilt stiffness of each pin 33 may be inthe range from 1.4×10¹⁶ to 4.7×10¹⁷ N² m²rad⁻².

FIG. 26 illustrates how the stiffnesses defined herein may be measured.FIG. 26 shows a plot of the displacement 6 resulting from theapplication of a load L (e.g. a force, moment or torque) applied to acomponent for which the stiffness is being measured. At levels of loadfrom zero to L_(P) there is a non-linear region in which displacement iscaused by motion of the component (or relative motion of separate partsof the component) as it is loaded, rather than deformation of thecomponent; for example moving within clearance between parts. At levelsof load above L_(Q) the elastic limit of the component has been exceededand the applied load no longer causes elastic deformation—plasticdeformation or failure of the component may occur instead. Betweenpoints P and Q the applied load and resulting displacement have a linearrelationship. The stiffnesses defined herein may be determined bymeasuring the gradient of the linear region between points P and Q (withthe stiffness being the inverse of that gradient). The gradient may befound for as large a region of the linear region as possible to increasethe accuracy of the measurement by providing a larger displacement tomeasure. For example, the gradient may be found by applying a load equalto or just greater than L_(P) and equal to or just less than L_(Q).Values for L_(P) and L_(Q) may be estimated prior to testing based onmaterials characteristics so as to apply suitable loads. Although thedisplacement is referred to as δ in this description, the skilled personwould appreciate that equivalent principles would apply to a linear orangular displacement.

The stiffnesses defined herein, unless otherwise stated, are for thecorresponding component(s) when the engine is off (i.e. at zero speed/onthe bench). The stiffnesses generally do not vary significantly over theoperating range of the engine; the stiffness at cruise conditions of theaircraft to which the engine is used (those cruise conditions being asdefined elsewhere herein) may therefore be the same as for when theengine is not in use. However, where the stiffness varies over theoperating range of the engine, the stiffnesses defined herein are to beunderstood as being values for when the engine is at room temperatureand unmoving.

The present disclosure also relates to methods 1300 of operating a gasturbine engine 10 on an aircraft. The methods 1300 are illustrated inFIG. 27. The method 1300 comprises starting up and operating 1302 theengine 10 (e.g taxiing on a runway, take-off, and climb of the aircraft,as suitable) to reach cruise conditions.

Once cruise conditions have been reached, the method 1300 then comprisesoperating 1304 the gas turbine engine 10, which may be as described inone or more embodiments elsewhere herein, to provide propulsion undercruise conditions.

The gas turbine engine 10 is such that, and/or is operated such that,any or all of the parameters or ratios defined herein are within thespecified ranges.

The torque on the core shaft 26 may be referred to as the input torque,as this is the torque which is input to the gearbox 30. Torque has unitsof [force]x[distance] and may be expressed in units of Newton metres(N.m), and is defined in the usual way as would be understood by theskilled person.

The torque supplied by the turbine 19 to the core shaft (i.e. the torqueon the core shaft) at cruise conditions may be greater than or equal to10,000 Nm, and optionally greater than or equal to 11,000 Nm. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the torque on the core shaft 26 at cruiseconditions may be greater than or equal to 10,000 Nm (and optionally maybe equal to 12,760 Nm). In some embodiments, for example in embodimentsin which the fan diameter is in the range from 330 to 380 cm, the torqueon the core shaft 26 at cruise conditions may be greater than or equalto 25,000 Nm, and optionally greater than or equal to 30,000 Nm (andoptionally may be equal to 33,970 Nm, or 34,000 Nm).

The torque on the core shaft at cruise conditions may be in the rangefrom 10,000 to 50,000 Nm, and optionally from 11,000 to 45,000 Nm. Insome embodiments, for example in embodiments in which the fan diameteris in the range from 240 to 280 cm, the torque on the core shaft 26 atcruise conditions may be in the range from 10,000 to 15,000 Nm, andoptionally from 11,000 to 14,000 Nm (and optionally may be equal to12,760 Nm). In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the torque on the coreshaft 26 at cruise conditions may be in the range from 25,000 Nm or50,000 Nm, and optionally from 30,000 to 40,000 Nm (and optionally maybe equal to 33,970 Nm, or 34,000 Nm).

Under maximum take-off (MTO) conditions, the torque on the core shaft 26may be greater than or equal to 28,000 Nm, and optionally greater thanor equal to 30,000 Nm. In some embodiments, for example in embodimentsin which the fan diameter is in the range from 240 to 280 cm, the torqueon the core shaft 26 under MTO conditions may be greater than or equalto 28,000, and optionally greater than or equal to 35,000 Nm (andoptionally may be equal to 36,300 Nm). In some embodiments, for examplein embodiments in which the fan diameter is in the range from 330 to 380cm, the torque on the core shaft 26 under MTO conditions may greaterthan or equal to 70,000 Nm, and optionally greater than or equal to80,000 or 82,000 Nm (and optionally may be equal to 87,000 Nm or 87,100Nm).

Under maximum take-off (MTO) conditions, the torque on the core shaft 26may be in the range from 28,000 Nm to 135,000 Nm, and optionally in therange from 30,000 to 110,000 Nm. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the torque on the core shaft 26 under MTO conditions may be in therange from 28,000 to 50,000 Nm, and optionally from 35,000 to 38,000 Nm(and optionally may be equal to 36,000 Nm or 36,300 Nm). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the torque on the core shaft 26 under MTOconditions may be in the range from 70,000 Nm or 135,000 Nm, andoptionally from 80,000 to 90,000 Nm or 82,000 to 92,000 Nm (andoptionally may be equal to 87,000 Nm or 87,100 Nm).

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

The invention claimed is:
 1. A gas turbine engine for an aircraftcomprising: an engine core comprising a turbine, a compressor, and acore shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; and a gearbox that receives an input from a gearbox input shaftportion of the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft, the gearboxbeing an epicyclic gearbox comprising a sun gear, a plurality of planetgears, a ring gear, and a planet carrier on which the planet gears aremounted, and wherein: a carrier to gearbox input shaft torsionalstiffness ratio of:$\frac{a\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{a\mspace{14mu}{torsional}{\mspace{11mu}\;}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{gearbox}\mspace{14mu}{input}\mspace{14mu}{shaft}}$is greater than or equal to
 70. 2. The gas turbine engine of claim 1wherein the carrier to gearbox input shaft torsional stiffness ratio isequal to or greater than
 75. 3. The gas turbine engine of claim 1wherein the carrier to gearbox input shaft torsional stiffness ratio isless than or equal to 5.0×10³.
 4. The gas turbine engine of claim 1wherein the carrier to gearbox input shaft torsional stiffness ratio isin a range from 7.5×10¹ to 3×10³.
 5. The gas turbine engine of claim 1wherein the torsional stiffness of the planet carrier is greater than orequal to 1.60×10⁸ Nm/rad, and in a range from 1.60×10⁸ to 1.00×10¹¹Nm/rad, or from 2.7×10⁸ to 1×10¹⁰ Nm/rad.
 6. The gas turbine engine ofclaim 1 wherein the torsional stiffness of the gearbox input shaft isequal to or greater than 1.4×10⁶ Nm/radian, and in a range from 1.4×10⁶to 2.5×10⁸ Nm/radian.
 7. The gas turbine engine of claim 1, wherein: (i)the fan has a fan diameter in a range from 240 to 280 cm and the carrierto gearbox input shaft torsional stiffness ratio is greater than orequal to 7.3×10¹; or (ii) the fan has a fan diameter in a range from 330to 380 cm and the carrier to gearbox input shaft torsional stiffnessratio is greater than or equal to 1.0×10².
 8. The gas turbine engine ofclaim 1, wherein: the turbine is a first turbine, the compressor is afirst compressor, and the core shaft is a first core shaft; the enginecore further comprises a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor; and the second turbine, second compressor, and second coreshaft are arranged to rotate at a higher rotational speed than the firstcore shaft.
 9. The gas turbine engine of claim 1, wherein the planetcarrier comprises a forward plate and a rearward plate and pinsextending therebetween, each pin being arranged to have one of theplurality of planet gears mounted thereon.
 10. The gas turbine engine ofclaim 9, wherein the planet carrier further comprises lugs extendingbetween the forward and rearward plates, the lugs being arranged to passbetween adjacent planet gears.
 11. The gas turbine engine of claim 1,wherein the gearbox comprises an odd number of planet gears, andcomprises 3, 5 or 7 planet gears.
 12. The gas turbine engine of claim 1,wherein the fan has a fan diameter greater than 240 cm and less than orequal to 380 cm, or greater than 300 cm and less than or equal to 380cm.
 13. The gas turbine engine of claim 1, wherein the gearbox is a stargearbox, in which the planet carrier does not rotate in use.
 14. The gasturbine engine of claim 1, wherein a pitch circle diameter of pins onwhich the planet gears are mounted is in a range from 0.38 to 0.65 m, orequal to 0.4 m or 0.55 m.
 15. The gas turbine engine of claim 1, whereinthe gearbox input shaft provides a soft mounting for the sun gear suchthat movement of the sun gear is facilitated.
 16. The gas turbine engineof claim 15, wherein the core shaft comprises a more stiff section and aless stiff section, the less stiff section providing the gearbox inputshaft and being arranged to lie between the more stiff section and thesun gear and arranged to provide, or to contribute to, the soft mountingof the sun gear.
 17. The gas turbine engine of claim 1, wherein: (i) agear ratio of the gearbox is in a range from 3.2 to 4.5, or from 3.3 to4.0; and/or (ii) a specific thrust of the engine at cruise is in a rangefrom 70 to 90 NKg⁻¹ s; and/or (iii) a bypass ratio at cruise is in arange from 12.5 to 18, or from 13 to
 16. 18. The gas turbine engine ofclaim 1, wherein a product of the torsional stiffness of the planetcarrier and the torsional stiffness of the gearbox input shaft isgreater than or equal to 1.5×10¹⁴ N²m²rad⁻², or greater than or equal to2.2×10¹⁴ N²m²rad⁻².
 19. The gas turbine engine of claim 1, wherein: (i)the fan has a fan diameter in a range from 240 to 280 cm and a productof the torsional stiffness of the planet carrier and the torsionalstiffness of the gearbox input shaft is greater than or equal to1.5×10¹⁴ N² m²rad⁻²; or (ii) the fan has a fan diameter in a range from330 to 380 cm and the product of the torsional stiffness of the planetcarrier and the torsional stiffness of the gearbox input shaft isgreater than or equal to 3.0×10¹⁵ N² m²rad⁻².
 20. A propulsor for anaircraft comprising: a fan comprising a plurality of fan blades; agearbox; and a power unit for driving the fan via the gearbox; whereinthe gearbox is an epicyclic gearbox arranged to receive an input from agearbox input shaft portion of a core shaft driven by the power unit andto output drive to the fan so as to drive the fan at a lower rotationalspeed than the core shaft, and comprises: a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier on which the planetgears are mounted, and wherein: a carrier to gearbox input shafttorsional stiffness ratio of:$\frac{a\mspace{14mu}{torsional}\mspace{14mu}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{planet}\mspace{14mu}{carrier}}{a\mspace{14mu}{torsional}{\mspace{11mu}\;}{stiffness}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{gearbox}\mspace{14mu}{input}\mspace{14mu}{shaft}}$is greater than or equal to 70.